Gearbox efficiency rating for turbomachine engines

ABSTRACT

A turbomachine engine can include a fan assembly, a vane assembly, a core engine, a gearbox, and a gearbox efficiency rating. The fan assembly can include a plurality of fan blades. The vane assembly can include a plurality of vanes, and the vanes can, in some instances, be disposed aft of the fan blades. The core engine can include one or more compressor sections and one or more turbine sections. The gearbox includes an input and an output. The input is coupled to the one or more turbine sections of the core engine and comprises a first rotational speed, the output is coupled to the fan assembly and has a second rotational speed, and a gear ratio of the first rotational speed to the second rotational speed is within a range of 4.1-14.0. The gearbox efficiency rating is 0.10-1.8.

ACKNOWLEDGMENT OF GOVERNMENT SUPPORT

The project leading to this application has received funding from theClean Sky 2 Joint Undertaking (JU) under grant agreement No. 945541. TheJU receives support from the European Union's Horizon 2020 research andinnovation programme and the Clean Sky 2 JU members other than theUnion.

CROSS-REFERENCE TO RELATED APPLICATION

This application claims the benefit of Italian Patent Application No.102020000019171, filed Aug. 4, 2020, which is incorporated by referenceherein.

FIELD

This disclosure relates generally to turbomachines including gearboxassemblies and, in particular, to apparatus and methods of determininggear assembly arrangements particular to certain turbomachineconfigurations.

BACKGROUND

A turbofan engine includes a core engine that drives a bypass fan. Thebypass fan generates the majority of the thrust of the turbofan engine.The generated thrust can be used to move a payload (e.g., an aircraft).

In some instances, a turbofan engine is configured as a direct driveengine. Direct drive engines are configured such that a power turbine(e.g., a low-pressure turbine) of the core engine is directly coupled tothe bypass fan. As such, the power turbine and the bypass fan rotate atthe same rotational speed (i.e., the same rpm).

In other instances, a turbofan engine can be configured as a gearedengine. Geared engines include a gearbox disposed between andinterconnecting the bypass fan and power turbine of the core engine. Thegearbox, for example, allows the power turbine of the core engine torotate at a different speed than the bypass fan. Thus, the gearbox can,for example, allow the power turbine of the core engine and the bypassfan to operate at their respective rotational speeds for maximumefficiency and/or power production.

Despite certain advantages, geared turbofan engines can have one or moredrawbacks. For example, including a gearbox in a turbofan engineintroduces additional complexity to the engine. This can, for example,make engine development and/or manufacturing significantly moredifficult. As such, there is a need for improved geared turbofanengines. There is also a need for devices and methods that can be usedto develop and manufacture geared turbofan engines more efficientlyand/or precisely.

BRIEF DESCRIPTION

Aspects and advantages of the disclosed technology will be set forth inpart in the following description, or may be obvious from thedescription, or may be learned through practice of the technologydisclosed in the description.

Various turbomachine engines and gear assemblies are disclosed herein.The disclosed turbomachine engines comprise a gearbox. And the disclosedturbomachine engines are characterized or defined by a gearboxefficiency rating. The gearbox efficiency rating (GER) equals

${Q\left( \frac{D^{1.56}}{T} \right)}^{1.53},$where Q is a gearbox oil flow rate an inlet of the gearbox measured ingallons per minute at a max takeoff condition, D is a diameter of thefan blades measured in inches, and T is a net thrust of the turbomachineengine measured in pounds force at the max takeoff condition. Thegearbox efficiency rating may also be used, for example, to aid thedevelopment of the gearbox in relation to other engine parameters. Thegearbox efficiency rating thus provides improved turbomachine enginesand/or can help simplify one or more complexities of geared turbomachineengine development.

In particular embodiments, a turbomachine engine includes a fanassembly, a vane assembly, a core engine, a gearbox, and a gearboxefficiency rating. The fan assembly includes a plurality of fan blades.The vane assembly includes a plurality of vanes. The core engineincludes one or more compressor sections and one or more turbinesections. The gearbox includes an input and an output. The input iscoupled to the one or more turbine sections of the core engine andcomprises a first rotational speed, the output is coupled to the fanassembly and has a second rotational speed, and a gear ratio of thefirst rotational speed to the second rotational speed is within a rangeof 4.1-14.0. The gearbox efficiency rating is 0.10-1.8.

These and other features, aspects, and/or advantages of the presentdisclosure will become better understood with reference to the followingdescription and the claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the disclosed technology and, together with thedescription, serve to explain the principles of the disclosure.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a cross-sectional schematic illustration of an exemplaryembodiment of a turbomachine engine configured with an open rotorpropulsion system.

FIG. 2 is a cross-sectional schematic illustration of an exemplaryembodiment of a turbomachine engine configured with an open rotorpropulsion system.

FIG. 3 is a cross-sectional schematic illustration of an exemplaryembodiment of a turbomachine engine configured with a ducted propulsionsystem.

FIG. 4 is a cross-sectional schematic illustration of an exemplaryembodiment of a turbomachine engine configured with a ducted propulsionsystem.

FIG. 5 is a cross-sectional schematic illustration of an exemplaryembodiment of a counter-rotating low-pressure turbine of a turbomachineengine, the low-pressure turbine having a 3×3 configuration.

FIG. 6 is a cross-sectional schematic illustration of an exemplaryembodiment of a counter-rotating low-pressure turbine of a turbomachineengine, the low-pressure turbine having a 4×3 configuration.

FIG. 7A is a graph depicting an exemplary range of gearbox efficiencyratings relative to an exemplary range of gear ratios for a turbomachineengine.

FIG. 7B is a graph depicting another exemplary range of gearboxefficiency ratings relative to an exemplary range of gear ratios for aturbomachine engine.

FIG. 7C is a graph depicting another exemplary range of gearboxefficiency ratings relative to an exemplary range of gear ratios for aturbomachine engine.

FIG. 8 is a chart depicting various engine parameters of severalexemplary turbomachine engines.

FIG. 9 is a cross-sectional schematic illustration of an exemplaryembodiment of a gearbox configuration for a turbomachine engine.

FIG. 10 is a cross-sectional schematic illustration of an exemplaryembodiment of a gearbox configuration for a turbomachine engine.

FIG. 11 is a cross-sectional schematic illustration of an exemplaryembodiment of a gearbox configuration for a turbomachine engine.

FIG. 12 is a cross-sectional schematic illustration of an exemplaryembodiment of a gearbox configuration for a turbomachine engine.

FIG. 13 is a cross-sectional schematic illustration of an exemplaryembodiment of a gearbox configuration for a turbomachine engine.

FIG. 14 is a schematic diagram of an exemplary lubricant systemsupplying lubricant to an engine component.

FIG. 15 is a schematic diagram of the lubricant system configured tosupply lubricant to a gearbox.

DETAILED DESCRIPTION

Reference now will be made in detail to embodiments of the disclosedtechnology, one or more examples of which are illustrated in thedrawings. Each example is provided by way of explanation of thedisclosed technology, not limitation of the disclosure. In fact, it willbe apparent to those skilled in the art that various modifications andvariations can be made in the present disclosure without departing fromthe scope or spirit of the disclosure. For instance, featuresillustrated or described as part of one embodiment can be used withanother embodiment to yield a still further embodiment. Thus, it isintended that the present disclosure covers such modifications andvariations as come within the scope of the appended claims and theirequivalents.

The word “exemplary” is used herein to mean “serving as an example,instance, or illustration.” Any implementation described herein as“exemplary” is not necessarily to be construed as preferred oradvantageous over other implementations.

As used herein, the terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.

The terms “forward” and “aft” refer to relative positions within a gasturbine engine or vehicle, and refer to the normal operational attitudeof the gas turbine engine or vehicle. For example, with regard to a gasturbine engine, forward refers to a position closer to an engine inletand aft refers to a position closer to an engine nozzle or exhaust.

The terms “upstream” and “downstream” refer to the relative directionwith respect to fluid flow in a fluid pathway. For example, “upstream”refers to the direction from which the fluid flows, and “downstream”refers to the direction to which the fluid flows.

The terms “coupled,” “fixed,” “attached to,” and the like refer to bothdirect coupling, fixing, or attaching, as well as indirect coupling,fixing, or attaching through one or more intermediate components orfeatures, unless otherwise specified herein.

The singular forms “a”, “an”, and “the” include plural references unlessthe context clearly dictates otherwise.

Approximating language, as used herein throughout the specification andclaims, is applied to modify any quantitative representation that couldpermissibly vary without resulting in a change in the basic function towhich it is related. Accordingly, a value modified by a term or terms,such as “about,” “approximately,” and “substantially,” are not to belimited to the precise value specified. In at least some instances, theapproximating language may correspond to the precision of an instrumentfor measuring the value, or the precision of the methods or machines forconstructing or manufacturing the components and/or systems. Forexample, the approximating language may refer to being within a 1, 2, 4,5, 10, 15, or 20 percent margin in either individual values, range(s) ofvalues and/or endpoints defining range(s) of values.

Here and throughout the specification and claims, range limitations arecombined and interchanged, such ranges are identified and include allthe sub-ranges contained therein unless context or language indicatesotherwise. For example, all ranges disclosed herein are inclusive of theendpoints, and the endpoints are independently combinable with eachother.

One or more components of the turbomachine engine or gear assemblydescribed hereinbelow may be manufactured or formed using any suitableprocess, such as an additive manufacturing process, such as a 3-Dprinting process. The use of such a process may allow such component tobe formed integrally, as a single monolithic component, or as anysuitable number of sub-components. In particular, the additivemanufacturing process may allow such component to be integrally formedand include a variety of features not possible when using priormanufacturing methods. For example, the additive manufacturing methodsdescribed herein enable the manufacture of heat exchangers having uniquefeatures, configurations, thicknesses, materials, densities, fluidpassageways, headers, and mounting structures that may not have beenpossible or practical using prior manufacturing methods. Some of thesefeatures are described herein.

Referring now to the drawings, FIG. 1 is an exemplary embodiment of anengine 100 including a gear assembly 102 according to aspects of thepresent disclosure. The engine 100 includes a fan assembly 104 driven bya core engine 106. In various embodiments, the core engine 106 is aBrayton cycle system configured to drive the fan assembly 104. The coreengine 106 is shrouded, at least in part, by an outer casing 114. Thefan assembly 104 includes a plurality of fan blades 108. A vane assembly110 extends from the outer casing 114 in a cantilevered manner. Thus,the vane assembly 110 can also be referred to as an unducted vaneassembly. The vane assembly 110, including a plurality of vanes 112, ispositioned in operable arrangement with the fan blades 108 to providethrust, control thrust vector, abate or re-direct undesired acousticnoise, and/or otherwise desirably alter a flow of air relative to thefan blades 108.

In some embodiments, the fan assembly 104 includes eight (8) to twenty(20) fan blades 108. In particular embodiments, the fan assembly 104includes ten (10) to eighteen (18) fan blades 108. In certainembodiments, the fan assembly 104 includes twelve (12) to sixteen (16)fan blades 108. In some embodiments, the vane assembly 110 includesthree (3) to thirty (30) vanes 112. In certain embodiments, the vaneassembly 110 includes an equal or fewer quantity of vanes 112 to fanblades 108. For example, in particular embodiments, the engine 100includes twelve (12) fan blades 108 and ten (10) vanes 112. In otherembodiments, the vane assembly 110 includes a greater quantity of vanes112 to fan blades 108. For example, in particular embodiments, theengine 100 includes ten (10) fan blades 108 and twenty-three (23) vanes112.

In certain embodiments, such as depicted in FIG. 1, the vane assembly110 is positioned downstream or aft of the fan assembly 104. However, itshould be appreciated that in some embodiments, the vane assembly 110may be positioned upstream or forward of the fan assembly 104. In stillvarious embodiments, the engine 100 may include a first vane assemblypositioned forward of the fan assembly 104 and a second vane assemblypositioned aft of the fan assembly 104. The fan assembly 104 may beconfigured to desirably adjust pitch at one or more fan blades 108, suchas to control thrust vector, abate or re-direct noise, and/or alterthrust output. The vane assembly 110 may be configured to desirablyadjust pitch at one or more vanes 112, such as to control thrust vector,abate or re-direct noise, and/or alter thrust output. Pitch controlmechanisms at one or both of the fan assembly 104 or the vane assembly110 may co-operate to produce one or more desired effects describedabove.

In certain embodiments, such as depicted in FIG. 1, the engine 100 is anun-ducted thrust producing system, such that the plurality of fan blades108 is unshrouded by a nacelle or fan casing. As such, in variousembodiments, the engine 100 may be configured as an unshrouded turbofanengine, an open rotor engine, or a propfan engine. In particularembodiments, the engine 100 is an unducted rotor engine with a singlerow of fan blades 108. The fan blades 108 can have a large diameter,such as may be suitable for high bypass ratios, high cruise speeds(e.g., comparable to aircraft with turbofan engines, or generally highercruise speed than aircraft with turboprop engines), high cruise altitude(e.g., comparable to aircraft with turbofan engines, or generally highercruise speed than aircraft with turboprop engines), and/or relativelylow rotational speeds.

The fan blades 108 comprise a diameter (D_(fan)). It should be notedthat for purposes of illustration only half of the D_(fan) is shown(i.e., the radius of the fan). In some embodiments, the D_(fan) is72-216 inches. In particular embodiments the D_(fan) is 100-200 inches.In certain embodiments, the D_(fan) is 120-190 inches. In otherembodiments, the D_(fan) is 72-120 inches. In yet other embodiments, theD_(fan) is 50-80 inches.

In some embodiments, the fan blade tip speed at a cruise flightcondition can be 650 to 900 fps, or 700 to 800 fps. A fan pressure ratio(FPR) for the fan assembly 104 can be 1.04 to 1.10, or in someembodiments 1.05 to 1.08, as measured across the fan blades at a cruiseflight condition.

Cruise altitude is generally an altitude at which an aircraft levelsafter climb and prior to descending to an approach flight phase. Invarious embodiments, the engine is applied to a vehicle with a cruisealtitude up to approximately 65,000 ft. In certain embodiments, cruisealtitude is between approximately 28,000 ft. and approximately 45,000ft. In still certain embodiments, cruise altitude is expressed in flightlevels (FL) based on a standard air pressure at sea level, in which acruise flight condition is between FL280 and FL650. In anotherembodiment, cruise flight condition is between FL280 and FL450. In stillcertain embodiments, cruise altitude is defined based at least on abarometric pressure, in which cruise altitude is between approximately4.85 psia and approximately 0.82 psia based on a sea-level pressure ofapproximately 14.70 psia and sea-level temperature at approximately 59degrees Fahrenheit. In another embodiment, cruise altitude is betweenapproximately 4.85 psia and approximately 2.14 psia. It should beappreciated that in certain embodiments, the ranges of cruise altitudedefined by pressure may be adjusted based on a different referencesea-level pressure and/or sea-level temperature.

The core engine 106 is generally encased in outer casing 114 definingone half of a core diameter (D_(core)), which may be thought of as themaximum extent from the centerline axis (datum for R). In certainembodiments, the engine 100 includes a length (L) from a longitudinally(or axial) forward end 116 to a longitudinally aft end 118. In variousembodiments, the engine 100 defines a ratio of L/D_(core) that providesfor reduced installed drag. In one embodiment, L/D_(core) is at least 2.In another embodiment, L/D_(core) is at least 2.5. In some embodiments,the L/D_(core), is less than 5, less than 4, and less than 3. In variousembodiments, it should be appreciated that the L/D_(core) is for asingle unducted rotor engine.

The reduced installed drag may further provide for improved efficiency,such as improved specific fuel consumption. Additionally, oralternatively, the reduced drag may provide for cruise altitude engineand aircraft operation at or above Mach 0.5. In certain embodiments, theL/D_(core), the fan assembly 104, and/or the vane assembly 110separately or together configure, at least in part, the engine 100 tooperate at a maximum cruise altitude operating speed betweenapproximately Mach 0.55 and approximately Mach 0.85; or betweenapproximately 0.72 to 0.85 or between approximately 0.75 to 0.85.

Referring still to FIG. 1, the core engine 106 extends in a radialdirection (R) relative to an engine centerline axis 120. The gearassembly 102 receives power or torque from the core engine 106 through apower input source 122 and provides power or torque to drive the fanassembly 104, in a circumferential direction C about the enginecenterline axis 120, through a power output source 124.

The gear assembly 102 of the engine 100 can include a plurality ofgears, including an input and an output. The gear assembly can alsoinclude one or more intermediate gears disposed between and/orinterconnecting the input and the output. The input can be coupled to aturbine section of the core engine 106 and can comprise a firstrotational speed. The output can be coupled to the fan assembly and canhave a second rotational speed. In some embodiments, a gear ratio of thefirst rotational speed to the second rotational speed is greater than4.1 (e.g., within a range of 4.1-14.0).

The gear assembly 102 (which can also be referred to as “a gearbox”) cancomprise various types and/or configuration. For example, in someembodiments, the gearbox is an epicyclic gearbox configured in a stargear configuration. Star gear configurations comprise a sun gear, aplurality of star gears (which can also be referred to as “planetgears”), and a ring gear. The sun gear is the input and is coupled tothe power turbine (e.g., the low-pressure turbine) such that the sungear and the power turbine rotate at the same rotational speed. The stargears are disposed between and interconnect the sun gear and the ringgear. The star gears are rotatably coupled to a fixed carrier. As such,the star gears can rotate about their respective axes but cannotcollectively orbit relative to the sun gear or the ring gear. As anotherexample, the gearbox is an epicyclic gearbox configured in a planet gearconfiguration. Planet gear configurations comprise a sun gear, aplurality of planet gears, and a ring gear. The sun gear is the inputand is coupled to the power turbine. The planet gears are disposedbetween and interconnect the sun gear and the ring gear. The planetgears are rotatably coupled to a rotatable carrier. As such, the planetgears can rotate about their respective axes and also collectivelyrotate together with the carrier relative to the sun gear and the ringgear. The carrier is the output and is coupled to the fan assembly. Thering gear is fixed from rotation.

In some embodiments, the gearbox is a single-stage gearbox (e.g., FIGS.10-11). In other embodiments, the gearbox is a multi-stage gearbox(e.g., FIGS. 9 and 12). In some embodiments, the gearbox is an epicyclicgearbox. In some embodiments, the gearbox is a non-epicyclic gearbox(e.g., a compound gearbox—FIG. 13).

As noted above, the gear assembly can be used to reduce the rotationalspeed of the output relative to the input. In some embodiments, a gearratio of the input rotational speed to the output rotational speed isgreater than 4.1. For example, in particular embodiments, the gear ratiois within a range of 4.1-14.0, within a range of 4.5-14.0, or within arange of 6.0-14.0. In certain embodiments, the gear ratio is within arange of 4.5-12 or within a range of 6.0-11.0. As such, in someembodiments, the fan assembly can be configured to rotate at arotational speed of 700-1500 rpm at a cruise flight condition, while thepower turbine (e.g., the low-pressure turbine) is configured to rotateat a rotational speed of 2,500-15,000 rpm at a cruise flight condition.In particular embodiments, the fan assembly can be configured to rotateat a rotational speed of 850-1350 rpm at a cruise flight condition,while the power turbine is configured to rotate at a rotational speed of5,000-10,000 rpm at a cruise flight condition.

Various gear assembly configurations are depicted schematically in FIGS.9-13. These gearboxes can be used any of the engines disclosed herein,including the engine 100. Additional details regarding the gearboxes areprovided below.

FIG. 2 shows a cross-sectional view of an engine 200, which isconfigured as an exemplary embodiment of an open rotor propulsionengine. The engine 200 is generally similar to the engine 100 andcorresponding components have been numbered similarly. For example, thegear assembly of the engine 100 is numbered “102” and the gear assemblyof the engine 200 is numbered “202,” and so forth. In addition to thegear assembly 202, the engine 200 comprises a fan assembly 204 thatincludes a plurality of fan blades 208 distributed around the enginecenterline axis 220. Fan blades 208 are circumferentially arranged in anequally spaced relation around the engine centerline axis 220, and eachfan blade 208 has a root 225 and a tip 226, and an axial span definedtherebetween, as well as a central blade axis 228.

The core engine 206 includes a compressor section 230, a combustionsection 232, and a turbine section 234 (which may be referred to as “anexpansion section”) together in a serial flow arrangement. The coreengine 206 extends circumferentially relative to an engine centerlineaxis 220. The core engine 206 includes a high-speed spool that includesa high-speed compressor 236 and a high-speed turbine 238 operablyrotatably coupled together by a high-speed shaft 240. The combustionsection 232 is positioned between the high-speed compressor 236 and thehigh-speed turbine 238.

The combustion section 232 may be configured as a deflagrativecombustion section, a rotating detonation combustion section, a pulsedetonation combustion section, and/or other appropriate heat additionsystem. The combustion section 232 may be configured as one or more of arich-burn system or a lean-burn system, or combinations thereof. Instill various embodiments, the combustion section 232 includes anannular combustor, a can combustor, a cannular combustor, a trappedvortex combustor (TVC), or other appropriate combustion system, orcombinations thereof.

The core engine 206 also includes a booster or low-pressure compressorpositioned in flow relationship with the high-pressure compressor 236.The low-pressure compressor 242 is rotatably coupled with thelow-pressure turbine 244 via a low-speed shaft 246 to enable thelow-pressure turbine 244 to drive the low-pressure compressor 242. Thelow-speed shaft 246 is also operably connected to the gear assembly 202to provide power to the fan assembly 204, such as described furtherherein.

It should be appreciated that the terms “low” and “high,” or theirrespective comparative degrees (e.g., “lower” and “higher”, whereapplicable), when used with compressor, turbine, shaft, or spoolcomponents, each refer to relative pressures and/or relative speedswithin an engine unless otherwise specified. For example, a “low spool”or “low-speed shaft” defines a component configured to operate at arotational speed, such as a maximum allowable rotational speed, lowerthan a “high spool” or “high-speed shaft” of the engine. Alternatively,unless otherwise specified, the aforementioned terms may be understoodin their superlative degree. For example, a “low turbine” or “low-speedturbine” may refer to the lowest maximum rotational speed turbine withina turbine section, a “low compressor” or “low speed compressor” mayrefer to the lowest maximum rotational speed turbine within a compressorsection, a “high turbine” or “high-speed turbine” may refer to thehighest maximum rotational speed turbine within the turbine section, anda “high compressor” or “high-speed compressor” may refer to the highestmaximum rotational speed compressor within the compressor section.Similarly, the low-speed spool refers to a lower maximum rotationalspeed than the high-speed spool. It should further be appreciated thatthe terms “low” or “high” in such aforementioned regards mayadditionally, or alternatively, be understood as relative to minimumallowable speeds, or minimum or maximum allowable speeds relative tonormal, desired, steady state, etc. operation of the engine.

The compressors and/or turbines disclosed herein can include variousstage counts. As disclosed herein the stage count includes the number ofrotors or blade stages in a particular component (e.g., a compressor orturbine). For example, in some embodiments, a low-pressure compressorcan comprise 1-8 stages, a high-pressure compressor can comprise 8-15stages, a high-pressure turbine comprises 1-2 stages, and/or alow-pressure turbine comprises 3-7 stages. For example, in certainembodiments, an engine can comprise a one stage low-pressure compressor,an 11 stage high-pressure compressor, a two stage high-pressurecompressor, and a 7 stage low-pressure turbine. As another example, anengine can comprise a three stage low-pressure compressor, a 10 stagehigh-pressure compressor, a two stage high-pressure compressor, and a 7stage low-pressure turbine.

In some embodiments, a low-pressure turbine is a counter-rotatinglow-pressure turbine comprising inner blade stages and outer bladestages. The inner blade stages extend radially outwardly from an innershaft, and the outer blade stages extend radially inwardly from an outerdrum. In particular embodiments, the counter-rotating low-pressureturbine comprises three inner blade stages and three outer blade stages,which can collectively be referred to as a six stage low-pressureturbine. In other embodiments, the counter-rotating low-pressure turbinecomprises four inner blade stages and three outer blade stages, whichcan be collectively be referred to as a seven stage low-pressureturbine.

As discussed in more detail below, the core engine 206 includes the gearassembly 202 that is configured to transfer power from the turbinesection 234 and reduce an output rotational speed at the fan assembly204 relative to the low-speed turbine 244. Embodiments of the gearassembly 202 depicted and described herein can allow for gear ratiossuitable for large-diameter unducted fans (e.g., gear ratios of4.1-14.0, 4.5-14.0, and/or 6.0-14.0). Additionally, embodiments of thegear assembly 202 provided herein may be suitable within the radial ordiametrical constraints of the core engine 206 within the outer casing214.

Various gearbox configurations are depicted schematically in FIGS. 9-13.These gearboxes can be used in any of the engines disclosed herein,including the engine 200. Additional details regarding the gearboxes areprovided below.

Engine 200 also includes a vane assembly 210 comprising a plurality ofvanes 212 disposed around engine centerline axis 220. Each vane 212 hasa root 248 and a tip 250, and a span defined therebetween. Vanes 212 canbe arranged in a variety of manners. In some embodiments, for example,they are not all equidistant from the rotating assembly.

In some embodiments, vanes 212 are mounted to a stationary frame and donot rotate relative to the engine centerline axis 220, but may include amechanism for adjusting their orientation relative to their axis 254and/or relative to the fan blades 208. For reference purposes, FIG. 2depicts a forward direction denoted with arrow F, which in turn definesthe forward and aft portions of the system.

As depicted in FIG. 2, the fan assembly 204 is located forward of thecore engine 106 with the exhaust 256 located aft of core engine 206 in a“puller” configuration. Other configurations are possible andcontemplated as within the scope of the present disclosure, such as whatmay be termed a “pusher” configuration embodiment where the engine coreis located forward of the fan assembly. The selection of “puller” or“pusher” configurations may be made in concert with the selection ofmounting orientations with respect to the airframe of the intendedaircraft application, and some may be structurally or operationallyadvantageous depending upon whether the mounting location andorientation are wing-mounted, fuselage-mounted, or tail-mountedconfigurations.

Left- or right-handed engine configurations, useful for certaininstallations in reducing the impact of multi-engine torque upon anaircraft, can be achieved by mirroring the airfoils (e.g., 208, 212)such that the fan assembly 204 rotates clockwise for one propulsionsystem and counterclockwise for the other propulsion system.Alternatively, an optional reversing gearbox can be provided to permit acommon gas turbine core and low-pressure turbine to be used to rotatethe fan blades either clockwise or counterclockwise, i.e., to provideeither left- or right-handed configurations, as desired, such as toprovide a pair of oppositely-rotating engine assemblies can be providedfor certain aircraft installations while eliminating the need to haveinternal engine parts designed for opposite rotation directions.

The engine 200 also includes the gear assembly 202 which includes a gearset for decreasing the rotational speed of the fan assembly 204 relativeto the low-speed turbine 244. In operation, the rotating fan blades 208are driven by the low-speed turbine 244 via gear assembly 202 such thatthe fan blades 208 rotate around the engine centerline axis 220 andgenerate thrust to propel the engine 200, and hence an aircraft on whichit is mounted, in the forward direction F.

In some embodiments, a gear ratio of the input rotational speed to theoutput rotational speed is greater than 4.1. For example, in particularembodiments, the gear ratio is within a range of 4.1-14.0, within arange of 4.5-14.0, or within a range of 6.0-14.0. In certainembodiments, the gear ratio is within a range of 4.5-12 or within arange of 6.0-11.0. As such, in some embodiments, the fan assembly can beconfigured to rotate at a rotational speed of 700-1500 rpm at a cruiseflight condition, while the power turbine (e.g., the low-pressureturbine) is configured to rotate at a rotational speed of 5,000-10,000rpm at a cruise flight condition. In particular embodiments, the fanassembly can be configured to rotate at a rotational speed of 850-1350rpm at a cruise flight condition, while the power turbine is configuredto rotate at a rotational speed of 5,500-9,500 rpm a cruise flightcondition.

It may be desirable that either or both of the fan blades 208 or thevanes 212 to incorporate a pitch change mechanism such that the bladescan be rotated with respect to an axis of pitch rotation (annotated as228 and 254, respectively) either independently or in conjunction withone another. Such pitch change can be utilized to vary thrust and/orswirl effects under various operating conditions, including to provide athrust reversing feature which may be useful in certain operatingconditions such as upon landing an aircraft.

Vanes 212 can be sized, shaped, and configured to impart a counteractingswirl to the fluid so that in a downstream direction aft of both fanblades 208 and vanes 212 the fluid has a greatly reduced degree ofswirl, which translates to an increased level of induced efficiency.Vanes 212 may have a shorter span than fan blades 208, as shown in FIG.2. For example, vanes 212 may have a span that is at least 50% of a spanof fan blades 208. In some embodiments, the span of the vanes can be thesame or longer than the span as fan blades 208, if desired. Vanes 212may be attached to an aircraft structure associated with the engine 200,as shown in FIG. 2, or another aircraft structure such as a wing, pylon,or fuselage. Vanes 212 may be fewer or greater in number than, or thesame in number as, the number of fan blades 208. In some embodiments,the number of vanes 212 are greater than two, or greater than four, innumber. Fan blades 208 may be sized, shaped, and contoured with thedesired blade loading in mind.

In the embodiment shown in FIG. 2, an annular 360-degree inlet 258 islocated between the fan assembly 204 and the vane assembly 210, andprovides a path for incoming atmospheric air to enter the core engine206 radially inwardly of at least a portion of the vane assembly 210.Such a location may be advantageous for a variety of reasons, includingmanagement of icing performance as well as protecting the inlet 258 fromvarious objects and materials as may be encountered in operation.

In the exemplary embodiment of FIG. 2, in addition to the open rotor orunducted fan assembly 204 with its plurality of fan blades 208, anoptional ducted fan assembly 260 is included behind fan assembly 204,such that the engine 200 includes both a ducted and an unducted fanwhich both serve to generate thrust through the movement of air atatmospheric temperature without passage through the core engine 206. Theducted fan assembly 260 is shown at about the same axial location as thevane 212, and radially inward of the root 248 of the vane 212.Alternatively, the ducted fan assembly 260 may be between the vane 212and core duct 262, or be farther forward of the vane 212. The ducted fanassembly 260 may be driven by the low-pressure turbine 244, or by anyother suitable source of rotation, and may serve as the first stage ofthe low-pressure compressor 242 or may be operated separately. Airentering the inlet 258 flows through an inlet duct 264 and then isdivided such that a portion flows through a core duct 262 and a portionflows through a fan duct 266. Fan duct 266 may incorporate heatexchangers 268 and exhausts to the atmosphere through an independentfixed or variable nozzle 270 aft of the vane assembly 210, at the aftend of the fan cowl 252 and outside of the engine core cowl 272. Airflowing through the fan duct 266 thus “bypasses” the core of the engineand does not pass through the core.

Thus, in the exemplary embodiment, engine 200 includes an unducted fanformed by the fan blades 208, followed by the ducted fan assembly 260,which directs airflow into two concentric or non-concentric ducts 262and 266, thereby forming a three-stream engine architecture with threepaths for air which passes through the fan assembly 204.

In the exemplary embodiment shown in FIG. 2, a slidable, moveable,and/or translatable plug nozzle 274 with an actuator may be included inorder to vary the exit area of the nozzle 270. A plug nozzle istypically an annular, symmetrical device which regulates the open areaof an exit such as a fan stream or core stream by axial movement of thenozzle such that the gap between the nozzle surface and a stationarystructure, such as adjacent walls of a duct, varies in a scheduledfashion thereby reducing or increasing a space for airflow through theduct. Other suitable nozzle designs may be employed as well, includingthose incorporating thrust reversing functionality. Such an adjustable,moveable nozzle may be designed to operate in concert with other systemssuch as VBV's, VSV's, or blade pitch mechanisms and may be designed withfailure modes such as fully-open, fully-closed, or intermediatepositions, so that the nozzle 270 has a consistent “home” position towhich it returns in the event of any system failure, which may preventcommands from reaching the nozzle 270 and/or its actuator.

In some embodiments, a mixing device 276 can be included in a region aftof a core nozzle 278 to aid in mixing the fan stream and the core streamto improve acoustic performance by directing core stream outward and fanstream inward.

Since the engine 200 shown in FIG. 2 includes both an open rotor fanassembly 204 and a ducted fan assembly 260, the thrust output of bothand the work split between them can be tailored to achieve specificthrust, fuel burn, thermal management, and/or acoustic signatureobjectives which may be superior to those of a typical ducted fan gasturbine propulsion assembly of comparable thrust class. The ducted fanassembly 260, by lessening the proportion of the thrust required to beprovided by the unducted fan assembly 104, may permit a reduction in theoverall fan diameter of the unducted fan assembly and thereby providefor installation flexibility and reduced weight.

Operationally, the engine 200 may include a control system that managesthe loading of the respective open and ducted fans, as well aspotentially the exit area of the variable fan nozzle, to providedifferent thrust, noise, cooling capacity and other performancecharacteristics for various portions of the flight envelope and variousoperational conditions associated with aircraft operation. For example,in climb mode the ducted fan may operate at maximum pressure ratiothere-by maximizing the thrust capability of stream, while in cruisemode, the ducted fan may operate a lower pressure ratio, raising overallefficiency through reliance on thrust from the unducted fan. Nozzleactuation modulates the ducted fan operating line and overall engine fanpressure ratio independent of total engine airflow.

The ducted fan stream flowing through fan duct 266 may include one ormore heat exchangers 268 for removing heat from various fluids used inengine operation (such as an air-cooled oil cooler (ACOC), cooledcooling air (CCA), etc.). The heat exchangers 268 may take advantage ofthe integration into the fan duct 266 with reduced performance penalties(such as fuel efficiency and thrust) compared with traditional ductedfan architectures, due to not impacting the primary source of thrustwhich is, in this case, the unducted fan stream. Heat exchangers maycool fluids such as gearbox oil, engine sump oil, thermal transportfluids such as supercritical fluids or commercially availablesingle-phase or two-phase fluids (supercritical CO2, EGV, Slither 800,liquid metals, etc.), engine bleed air, etc. Heat exchangers may also bemade up of different segments or passages that cool different workingfluids, such as an ACOC paired with a fuel cooler. Heat exchangers 268may be incorporated into a thermal management system which provides forthermal transport via a heat exchange fluid flowing through a network toremove heat from a source and transport it to a heat exchanger.

Since the fan pressure ratio is higher for the ducted fan than for theunducted fan, the fan duct provides an environment where more compactheat exchangers may be utilized than would be possible if installed onthe outside of the core cowl in the unducted fan stream. Fan bypass airis at a very low fan pressure ratio (FPR) (1.05 to 1.08), making itdifficult to drive air through heat exchangers. Without the availabilityof a fan duct as described herein, scoops or booster bleed air may berequired to provide cooling air to and through heat exchangers. A set ofparameters can be developed around heat exchangers in the fan duct,based on heat load, heat exchanger size, ducted fan stream correctedflow, and ducted fan stream temperature.

The fan duct 266 also provides other advantages in terms of reducednacelle drag, enabling a more aggressive nacelle close-out, improvedcore stream particle separation, and inclement weather operation. Byexhausting the fan duct flow over the core cowl, this aids in energizingthe boundary layer and enabling the option of a steeper nacelle closeout angle between the maximum dimension of the engine core cowl 272 andthe exhaust 256. The close-out angle is normally limited by air flowseparation, but boundary layer energization by air from the fan duct 266exhausting over the core cowl reduces air flow separation. This yields ashorter, lighter structure with less frictional surface drag.

The fan assembly and/or vane assembly can be shrouded or unshrouded (asshown in FIGS. 1 and 2). Although not shown, an optional annular shroudor duct can be coupled to the vane assembly 210 and located distallyfrom the engine centerline axis 220 relative to the vanes 212. Inaddition to the noise reduction benefit, the duct may provide improvedvibratory response and structural integrity of the vanes 212 by couplingthem into an assembly forming an annular ring or one or morecircumferential sectors, i.e., segments forming portions of an annularring linking two or more of the vanes 212. The duct may also allow thepitch of the vanes to be varied more easily. For example, FIGS. 3-4,discussed in more detail below, disclose embodiments in which both thefan assembly and vane assembly are shrouded.

Although depicted above as an unshrouded or open rotor engine in theembodiments depicted above, it should be appreciated that aspects of thedisclosure provided herein may be applied to shrouded or ducted engines,partially ducted engines, aft-fan engines, or other turbomachineconfigurations, including those for marine, industrial, oraero-propulsion systems. Certain aspects of the disclosure may beapplicable to turbofan, turboprop, or turboshaft engines. However, itshould be appreciated that certain aspects of the disclosure may addressissues that may be particular to unshrouded or open rotor engines, suchas, but not limited to, issues related to gear ratios, fan diameter, fanspeed, length (L) of the engine, maximum diameter of the core engine(Dcore) of the engine, L/Dcore of the engine, desired cruise altitude,and/or desired operating cruise speed, or combinations thereof.

FIG. 3 is a schematic cross-sectional view of a gas turbine engine inaccordance with an exemplary embodiment of the present disclosure. Moreparticularly, for the embodiment of FIG. 3, the gas turbine engine is ahigh-bypass turbofan jet engine 300, referred to herein as “turbofanengine 300.” As shown in FIG. 3, the turbofan engine 300 defines anaxial direction A (extending parallel to a longitudinal centerline 302provided for reference) and a radial direction R (extendingperpendicular to the axial direction A). In general, the turbofan 300includes a fan section 304 and a core engine 306 disposed downstreamfrom the fan section 304. The engine 300 also includes a gear assemblyor power gear box 336 having a plurality of gears for coupling a gasturbine shaft to a fan shaft. The position of the power gear box 336 isnot limited to that as shown in the exemplary embodiment of turbofan300. For example, the position of the power gear box 336 may vary alongthe axial direction A.

The exemplary core engine 306 depicted generally includes asubstantially tubular outer casing 308 that defines an annular inlet310. The outer casing 308 encases, in serial flow relationship, acompressor section including a booster or low-pressure (LP) compressor312 and a high-pressure (HP) compressor 314; a combustion section 316; aturbine section including a high-pressure (HP) turbine 318 and alow-pressure (LP) turbine 320; and a jet exhaust nozzle section 322. Ahigh-pressure (HP) shaft or spool 324 drivingly connects the HP turbine318 to the HP compressor 314. A low-pressure (LP) shaft or spool 326drivingly connects the LP turbine 320 to the LP compressor 312.Additionally, the compressor section, combustion section 316, andturbine section together define at least in part a core air flowpath 327extending therethrough.

A gear assembly of the present disclosure is compatible with standardfans, variable pitch fans, or other configurations. For the embodimentdepicted, the fan section 304 may include a variable pitch fan 328having a plurality of fan blades 330 coupled to a disk 332 in a spacedapart manner. As depicted, the fan blades 330 extend outwardly from disk332 generally along the radial direction R. Each fan blade 330 isrotatable relative to the disk 332 about a pitch axis P by virtue of thefan blades 330 being operatively coupled to a suitable actuation member334 configured to collectively vary the pitch of the fan blades 330. Thefan blades 330, disk 332, and actuation member 334 are togetherrotatable about the longitudinal axis 302 by LP shaft 326 across a gearassembly or power gear box 336. A gear assembly 336 may enable a speedchange between a first shaft, e.g., LP shaft 326, and a second shaft,e.g., LP compressor shaft and/or fan shaft. For example, in oneembodiment, the gear assembly 336 may be disposed in an arrangementbetween a first shaft and a second shaft such as to reduce an outputspeed from one shaft to another shaft.

More generally, the gear assembly 336 can be placed anywhere along theaxial direction A to decouple the speed of two shafts, whenever it isconvenient to do so from a component efficiency point of view, e.g.,faster LP turbine and slower fan and LP compressor or faster LP turbineand LP compressor and slower fan.

Referring still to the exemplary embodiment of FIG. 3, the disk 332 iscovered by rotatable front nacelle 338 aerodynamically contoured topromote an airflow through the plurality of fan blades 330.Additionally, the exemplary fan section 304 includes an annular fancasing or outer nacelle 340 that circumferentially surrounds the fan 328and/or at least a portion of the core engine 306. The nacelle 340 is,for the embodiment depicted, supported relative to the core engine 306by a plurality of circumferentially-spaced outlet guide vanes 342.Additionally, a downstream section 344 of the nacelle 340 extends overan outer portion of the core engine 306 so as to define a bypass airflowpassage 346 therebetween.

During operation of the turbofan engine 300, a volume of air 348 entersthe turbofan 300 through an associated inlet 350 of the nacelle 340and/or fan section 304. As the volume of air 348 passes across the fanblades 330, a first portion of the air 348 as indicated by arrows 352 isdirected or routed into the bypass airflow passage 346 and a secondportion of the air 348 as indicated by arrow 354 is directed or routedinto the LP compressor 312. The ratio between the first portion of air352 and the second portion of air 354 is commonly known as a bypassratio. The pressure of the second portion of air 354 is then increasedas it is routed through the high-pressure (HP) compressor 314 and intothe combustion section 316, where it is mixed with fuel and burned toprovide combustion gases 356.

The combustion gases 356 are routed through the HP turbine 318 where aportion of thermal and/or kinetic energy from the combustion gases 356is extracted via sequential stages of HP turbine stator vanes 358 thatare coupled to the outer casing 308 and HP turbine rotor blades 360 thatare coupled to the HP shaft or spool 324, thus causing the HP shaft orspool 324 to rotate, thereby supporting operation of the HP compressor314. The combustion gases 356 are then routed through the LP turbine 320where a second portion of thermal and kinetic energy is extracted fromthe combustion gases 356 via sequential stages of LP turbine statorvanes 362 that are coupled to the outer casing 308 and LP turbine rotorblades 364 that are coupled to the LP shaft or spool 326, thus causingthe LP shaft or spool 326 to rotate, thereby supporting operation of theLP compressor 312 and/or rotation of the fan 328.

The combustion gases 356 are subsequently routed through the jet exhaustnozzle section 322 of the core engine 306 to provide propulsive thrust.Simultaneously, the pressure of the first portion of air 352 issubstantially increased as the first portion of air 352 is routedthrough the bypass airflow passage 346 before it is exhausted from a fannozzle exhaust section 366 of the turbofan 300, also providingpropulsive thrust. The HP turbine 318, the LP turbine 320, and the jetexhaust nozzle section 322 at least partially define a hot gas path 368for routing the combustion gases 356 through the core engine 306.

For example, FIG. 4 is a cross-sectional schematic illustration of anexemplary embodiment of an engine 400 that includes a gear assembly 402in combination with a ducted fan assembly 404 and a core engine 406.However, unlike the open rotor configuration of the engine 200, the fanassembly 404 and its fan blades 408 are contained within an annular fancase 480 (which can also be referred to as “a nacelle”) and the vaneassembly 410 and the vanes 412 extend radially between the fan cowl 452(and/or the engine core cowl 472) and the inner surface of the fan case480. As discussed above, the gear assemblies disclosed herein canprovide for increased gear ratios for a fixed gear envelope (e.g., withthe same size ring gear), or alternatively, a smaller diameter ring gearmay be used to achieve the same gear ratios.

The core engine 400 comprises a compressor section 430, a combustorsection 432, and a turbine section 434. The compressor section 430 caninclude a high-pressure compressor 436 and a booster or a low-pressurecompressor 442. The turbine section 434 can include a high-pressureturbine 438 and a low-pressure turbine 444. The low-pressure compressor442 is positioned forward of and in flow relationship with thehigh-pressure compressor 436. The low-pressure compressor 442 isrotatably coupled with the low-pressure turbine 444 via a low-speedshaft 446 to enable the low-pressure turbine 444 to drive thelow-pressure compressor 442 (and a ducted fan 460). The low-speed shaft446 is also operably connected to the gear assembly 402 to provide powerto the fan assembly 404. The high-pressure compressor 436 is rotatablycoupled with the high-pressure turbine 438 via a high-speed shaft 440 toenable the high-pressure turbine 438 to drive the high-pressurecompressor 436.

In some embodiments, the engine 400 can comprise a pitch changemechanism 482 coupled to the fan assembly 404 and configured to vary thepitch of the fan blades 408. In certain embodiments, the pitch changemechanism 482 can be a linear actuated pitch change mechanism.

In some embodiments, the engine 400 can comprise a variable fan nozzle.Operationally, the engine 400 may include a control system that managesthe loading of the fan, as well as potentially the exit area of thevariable fan nozzle, to provide different thrust, noise, coolingcapacity and other performance characteristics for various portions ofthe flight envelope and various operational conditions associated withaircraft operation. For example, nozzle actuation modulates the fanoperating line and overall engine fan pressure ratio independent oftotal engine airflow.

In some embodiments, an engine (e.g., the engine 100, the engine 200,and/or the engine 400) can comprise a counter-rotating low-pressureturbine. For example, FIGS. 5-6 depict schematic cross-sectionalillustrations of counter-rotating low-pressure turbines. In particular,FIG. 5 depicts a counter-rotating turbine 500, and FIG. 6 depicts acounter-rotating turbine 600. The counter-rotating turbines compriseinner blade stages and outer blade stages arranged in an alternatinginner-outer configuration. In other words, the counter-rotating turbinesdo not comprise stator vanes disposed between the blade stages.

Referring to FIG. 5, the counter-rotating turbine 500 comprises aplurality of inner blade stages 502 and a plurality of outer bladestages 504. More specifically, the counter-rotating turbine 500 includesthree inner blades stages 502 that are coupled to and extend radiallyoutwardly from an inner shaft 506 (which can also be referred to as “arotor”) and three outer blade stages 504 that are coupled to extendradially inwardly from an outer shaft 508 (which can also be referred toas “a drum”). In this manner, the counter-rotating turbine 500 can beconsidered a six stage turbine.

Referring to FIG. 6, the counter-rotating turbine 600 comprises aplurality of inner blade stages 602 and a plurality of outer bladestages 604. More specifically, the counter-rotating turbine 600 includesfour inner blades stages 602 that are coupled to and extend radiallyoutwardly from an inner shaft 606 and three outer blade stages 604 thatare coupled to extend radially inwardly from an outer shaft 608. In thismanner, the counter-rotating turbine 600 can be considered a seven stageturbine.

According to some embodiments there is a turbomachine characterized byboth a high gear ratio and a high power gearbox. A high gear ratiogearbox means a gearbox with a gear ratio of above about 4:1. Examplesof a high power gearbox include a gearbox adapted for transmitting powergreater than 7 MW with output spool speed above, e.g., 1000 rpm, agearbox adapted for transmitting power greater than 15 MW with outputspool speed of about 1100 rpm, and a gearbox adapted for transmittingpower greater than transmitting 22 MW with output spool speed of about3500 rpm.

Each of the embodiments of turbomachines disclosed herein utilize a highgear ratio gearbox. Adoption of a gearbox having a high gear ratiopresents unique challenges. One such challenge is determining the amountof oil that would need to circulate through the gearbox duringoperation, i.e., the high gear ratio gearbox's oil flow rate. The oildemand is significant when the engine requires a high gear ratiogearbox. Moreover, the estimated amount of oil flow for the high gearratio gearbox is not well informed by, or capable of being estimatedfrom, oil flow rates for an existing serviced engine. Starting from thisbasis, the inventors set out to calculate the oil flow demands for thedifferent engine embodiments contemplated and disclosed herein, byconsideration of the different features and performance characteristics,e.g., pitch line velocity and constants differentiating one gearboxconfiguration from another. The high gear ratio gearbox architecturesconsidered include those described and disclosed herein (e.g., FIGS.9-13 and the accompanying text, infra). These efforts accordinglyinvolved factoring in specific characteristics of the gearboxes and thepower transmission requirements for the gearbox to estimate the oil flowrates.

During the process of developing the aforementioned embodiments ofturbomachines incorporating a high gear ratio gearbox, the inventorsdiscovered, unexpectedly, that a good approximation of the high gearratio gearbox oil flow rate may be made using only a relatively fewengine parameters. This development is based on, among other things, therecognition that an oil flow rate through a gearbox is related to theexpected power loss when transmitting power across a gearbox. From thisinitial recognition and other developments that were the by-product ofstudying several different engine configurations that included a powergearbox (including the configurations disclosed herein), the inventorsultimately discovered that a good approximation to the high gear ratiogearbox oil flow rate could be made based on a relationship among theturbomachine's gearbox gear ratio, net thrust, and fan diameter. Theinventors refer to this relationship as a gearbox efficiency rating.

This discovery is quite beneficial. For example, with the gearboxefficiency rating having provided the engine oil flow requirements onecan also estimate, for purposes of system integration, the type ofoil-related secondary systems (e.g., sump, oil circuit, heat sinks,etc.) that would be included to support proper functioning of theselected high gear ratio gearbox; and/or to provide guidance on whethera particular engine architecture is beneficial or not, without requiringan entire team to complete the tedious and time-consuming process ofdeveloping a new gearbox from scratch. Therefore, the gearbox efficiencyrating can improve the process of developing a turbomachine engine,which can ultimately result in improved turbomachines.

As indicated, the gearbox efficiency rating is a relationship based on aturbomachine's fan diameter (D), net thrust (T), and gear ratio of ahigh gear ratio gearbox. The gear efficiency rating, valid for gearratios between about 4:1 and 14:1, may be expressed asQ(D^(1.56)/T)^(1.56), where Q is measured at an inlet of the gearbox ingallons per minute at a max takeoff condition, D is measured in inches,and T is measured in pounds force at the max takeoff condition. In thismanner, the gearbox efficiency rating defines a specific turbomachineengine configuration.

As used herein “net thrust” (T) equals the change of momementum of thebypass airflow plus the change of momentum of the core airflow and theburned fuel. Or stated another way, T=W_(byp)(V_(byp)−V₀)+(W_(core)+W_(fuel)) V_(core)−W_(core) V₀, where W_(byp) isthe mass flow rate of air of the bypass airflow, V_(byp) is the velocityof the bypass airflow, V₀ is the flight velocity, W_(core) is the massflow rate of air of the core airflow, W_(fuel) is the mass flow rate ofthe burned fuel, and V_(core) is the velocity of the core airflow.

As indicated earlier, turbomachine engines, such as the turbofan engines100, 200, 300, 400, comprise many variables and factors that affecttheir performance and/or operation. The interplay between the variouscomponents can make it particularly difficult to develop or select onecomponent, especially when each of the components is at a differentstage of completion. For example, one or more components may be nearlycomplete, yet one or more other components may be in an initial orpreliminary phase where only one (or a few) parameters is known. Also,each component is subject to change often more than once over thedevelopment period, which can often last for many years (e.g., 5-15years). These complex and intricate individual and collectivedevelopment processes can be cumbersome and inefficient. For at leastthese reasons, there is a need for devices and methods that can providea good estimate of, not only the basic configuration or sizing needed toachieve the desired performance benefits, but also to reflect thepenalties or accommodations in other areas in order to realize thedesired benefits.

According to another aspect of the disclosure, the gearbox efficiencyrating may additionally provide a particularly useful indication of theefficiency and effectiveness of the engine during initial development,e.g., as a tool to accept or reject a particular configuration. Thus,the gearbox efficiency rating can be used, for example, to guide gearboxdevelopment. For example, the gearbox efficiency rating can be used toquickly and accurately determine the size of the gearbox that issuitable for a particular engine without requiring an individual or teamto complete the tedious and time-consuming process of developing thegearbox from scratch. Therefore, the gearbox efficiency rating can alsoimprove the process of developing a turbomachine engine.

FIGS. 7A-8 illustrate exemplary ranges and/or values for gear efficiencyrating. FIGS. 7A-7C disclose exemplary ranges of gear efficiency ratingwith respect to various gear ratios. FIG. 8 discloses the gear ratio,oil flow, fan diameter, net thrust, and gearbox efficiency ratings formultiple exemplary turbomachine engines.

In some embodiments, the gearbox efficiency rating of a turbomachineengine is within a range of about 0.10-1.8 or 0.19-1.8 or 0.10-1.01. Incertain embodiments, the gearbox efficiency rating is within a range ofabout 0.25-0.55 or about 0.29-0.51. FIG. 8 provides the gear efficiencyrating of several exemplary engines.

In some embodiments, the oil flow rate Q is within a range of about 5-55gallons per minute. In certain embodiments, the oil flow rate Q iswithin a range of about 5.5-25 gallons per minute. FIG. 8 also providesthe oil flow rates of several exemplary engines.

As noted above, the oil flow rate Q is measured at an inlet of thegearbox in gallons per minute at a max takeoff condition. The inlet ofthe gearbox is the location at which the oil enters the gearbox from theoil supply line. As used herein “a max takeoff condition” meanssea-level elevation, standard pressure, extreme hot day temperature, anda flight velocity of up to about 0.25 Mach.

As used herein, the term “extreme hot day temperature” means the extremehot day temperature specified for a particular engine. This can includethe extreme hot day temperature used for engine certification. Extremehot day temperature can additionally or alternatively includetemperatures of about 130-140° F.

In some embodiments, the fan diameter D is about 120-216 inches. Incertain embodiments, the fan diameter D is about 120-192 inches. FIG. 8also provides the fan diameter of several exemplary engines.

In some embodiments, the net thrust T of the engine is within a range ofabout 10,000-100,000 pounds force. In particular embodiments, the netthrust T of the engine is within a range of about 12,000-30,000 poundsforce. FIG. 8 also provides the net thrust of several exemplary engines.

In some embodiments, the gearbox efficiency rating of a turbomachineengine can be configured in relation to the gear ratio (GR) of thegearbox. For example, in certain embodiments, a turbomachine engine canbe configured such that the gearbox efficiency rating is greater than0.015(GR^(1.4)) and less than 0.034(GR^(1.5)), as depicted in FIG. 7A.In other embodiments, a turbomachine engine can be configured such thatthe gearbox efficiency rating is greater than 0.02625(GR^(1.4)) and lessthan 0.042(GR^(1.4)).

For example, FIG. 8 depicts several examplary engines with gearboxefficiency ratings that satisfy these relationships. Engine 1 is aturbomachine engine comprising a gearbox with a gear ratio of 10.5:1 anda gearbox efficiency rating within a range of 0.40-1.16, specifically1.02. Engine 2, Engine 3, and Engine 4 are turbomachine enginescomprising gearboxes with a gear ratio of 7:1 and the gearbox efficiencyratings within a range of 0.23-0.63, that is 0.51, 0.42, and 0.41,respectively. Engine 5 is a turbomachine engine comprising a gearboxwith a gear ratio of 5.1:1 and a gearbox efficiency rating within arange of 0.15-0.39, specifically 0.29. Engine 6 is a turbomachine enginecomprising a gearbox with a gear ratio of 4.1:1 and a gearbox efficiencyrating within a range of 0.11-0.28, specifically 0.21. Engines 7-19provide additional examples with specific gearbox efficiency ratings.Ranges for the gearbox efficiency ratings of Engines 7-19 can bedetermined using the equations above and/or the charts of FIGS. 7A-7C.

As another example, a turbomachine engine comprising a gearbox with agear ratio of 4.5:1 can be configured such that the gearbox efficiencyrating is within a range of 0.12-0.32. As another example, aturbomachine engine comprising a gearbox with a gear ratio of 6:1 can beconfigured such that the gearbox efficiency rating within a range of0.18-0.50. As another example, a turbomachine engine comprising agearbox with a gear ratio of 9:1 can be configured such that the gearboxefficiency rating within a range of 0.33-0.92. As another example, aturbomachine engine comprising a gearbox with a gear ratio of 11:1 canbe configured such that the gearbox efficiency rating within a range of0.43-1.24. As another example, a turbomachine engine comprising agearbox with a gear ratio of 12:1 can be configured such that thegearbox efficiency rating within a range of 0.49-1.41. As yet anotherexample, a turbomachine engine comprising a gearbox with a gear ratio of14:1 can be configured such that the gearbox efficiency rating is withina range of 0.60-1.78.

In some instances, a turbomachine engine can comprise a gearbox with agear ratio of 5-6, 7-8, 9-10, 11-12, or 13-14. In other instances, aturbomachine engine can comprise a gearbox with a gear ratio of 5-7,8-10, 11-13. In yet other embodiments, a turbomachine engine cancomprise a gearbox with a gear ratio of 7-10 or 11-14. Below is a tablewith several exemplary gearbox efficiency ratings with respect toseveral exemplary gear ratios.

Gearbox Gear Efficiency Ratio Rating 4.1-6.9 0.10-0.62 7.0-9.9 0.22-1.0610.0-12.9 0.37-1.56 13.0-14.0 0.54-1.8 

In some embodiments, a turbomachine engine can be configured such thatthe gearbox efficiency rating is greater than 0.023 (GR^(1.5)) and lessthan 0.034(GR^(1.5)), as depicted in FIG. 7B. In particular instances,the gearbox efficiency rating can be about 0.0275(GR^(1.5)). Theseembodiments can be particularly advantageous, for example, with enginescomprising an epicyclic gearbox (e.g., star and/or planetconfiguration).

Gearbox Gear Efficiency Ratio Rating 4.1-6.9 0.19-0.62 7.0-9.9 0.43-1.0610.0-12.9 0.73-1.58 13.0-14.0 1.08-1.8 

In other embodiments, a turbomachine engine can be configured such thatthe gearbox efficiency rating is greater than 0.015(GR^(1.4)) and lessthan 0.025(GR^(1.4)), as depicted in FIG. 7C. In particular instances,the gearbox efficiency rating can be about 0.02 (GR^(1.4)). Theseembodiments can be particularly advantageous, for example, with enginescomprising a non-epicyclic gearbox (e.g., compound gearboxes).

Gearbox Gear Efficiency Ratio Rating 4.1-6.9 0.10-0.37 7.0-9.9 0.23-0.6210.0-12.9 0.38-0.90 13.0-14.0 0.54-1.01

It should be noted gearbox efficiency rating values disclosed herein areapproximate values. Accordingly, the disclosed gearbox efficiency ratingvalues include values within five percent of the listed values.

As noted above, the gearbox efficiency rating can define a specificengine configuration and/or can be used when developing a gearbox for aturbomachine engine. For example, in some instances, the gearboxefficiency rating can be used to determine the size and/or oil flow rateof a gearbox. Assuming that a desired gear ratio of the gearbox isknown, along with the fan diameter, and the net thrust of the engine,the gearbox efficiency ratings depicted in the charts of FIG. 7A-7C canbe used to determine an acceptable oil flow rate. In some embodiments,the equation below can be used to determine an acceptable range of oilflow rates (Q) for the gearbox. The determined oil flow rate Q can beused, for example, to aid in the configuration of the gearbox. In someinstances, one or more other parameters (e.g., the gearbox efficiencyrating) can also aid in the configuration of the gearbox.

$\frac{0.015\left( {GR}^{1.4} \right)}{\left( \frac{D^{1.56}}{T} \right)^{1.53}} < Q < \frac{0.034\left( {GR}^{1.5} \right)}{\left( \frac{D^{1.56}}{T} \right)^{1.53}}$

For example, a gearbox for a turbomachine engine can be configured usingthe following exemplary method. With reference to FIG. 8, Engine 1comprises an unducted fan and can be configured similar to the engine200. Engine 1 comprises a fan diameter of 188.6 inches and a net thrustof 25,503 pounds force at a max takeoff condition. Engine 1 furthercomprises a five stage low-pressure turbine. The desired gear ratio forthe gearbox of Engine 1 is about 10.5:1. Based on this information, oilflow rate Q of the gearbox of Engine 1 should be about 8-24 gallons perminute at a max takeoff condition.

FIG. 9 schematically depicts a gearbox 700 that can be used, forexample, with Engine 1. The gearbox 700 comprises a two-stage starconfiguration.

The first stage of the gearbox 700 includes a first-stage sun gear 702,a first-stage carrier 704 housing a plurality of first-stage star gears,and a first-stage ring gear 706. The first-stage sun gear 702 can becoupled to a low-speed shaft 708, which in turn is coupled to thelow-pressure turbine of Engine 1. The first-stage sun gear 702 can meshwith the first-stage star gears, which mesh with the first-stage ringgear. The first-stage carrier 704 can be fixed from rotation by asupport member 710.

The second stage of the gearbox 700 includes a second-stage sun gear712, a second-stage carrier 714 housing a plurality of second-stage stargears, and a second-stage ring gear 716. The second-stage sun gear 712can be coupled to a shaft 718 which in turn is coupled to thefirst-stage ring gear 706. The second-stage carrier 714 can be fixedfrom rotation by a support member 720. The second-stage ring gear 716can be coupled to a fan shaft 722.

In some embodiments, each stage of the gearbox 700 can comprise fivestar gears. In other embodiments, the gearbox 700 can comprise fewer ormore than five star gears in each stage. In some embodiments, thefirst-stage carrier can comprise a different number of star gears thanthe second-stage carrier. For example, the first-carrier can comprisefive star gears, and the second-stage carrier can comprise three stargears, or vice versa.

Based on the configuration of the gearbox 700 and the calculated oilflow rate of 8-24 gallons per minute, which is based on the gearboxefficiency rating, the gearbox 700 can comprise a radius R₁. The size ofthe gearbox, including the radius R₁, can be configured such that theoil flow rate at the inlet of the gearbox 700 at a max takeoff conditionis about 8-24 gallons per minute or about 16-24 gallons per minute(e.g., 20.9 gpm). In some embodiments, the radius R₁ of the gearbox 700can be about 16-19 inches. In other embodiments, the radius R₁ of thegearbox 700 can be about 22-24 inches. In other embodiments, the radiusR₁ of the gearbox 700 can be smaller than 16 inches or larger than 24inches.

As another example, Engine 2 (FIG. 8) comprises an unducted fan and canbe configured similar to the engine 200. Engine 3 comprises a fandiameter of 188.6 inches and a net thrust of 25,000 pounds force at amax takeoff condition. Engine 2 further comprises a 3-7 stagelow-pressure turbine. The desired gear ratio for the gearbox of Engine 2is about 7:1. Based on this information, oil flow rate Q of the gearboxof Engine 2 should be about 4-13 gallons per minute or about 8-13gallons per minute (e.g., 10.06) at a max takeoff condition.

FIG. 10 schematically depicts a gearbox 800 that can be used, forexample, with Engine 2. The gearbox 800 comprises a single-stage starconfiguration. The gearbox 800 includes a sun gear 802, a carrier 804housing a plurality of star gears (e.g., 3-5 star gears), and a ringgear 806. The sun gear 802 can mesh with the star gears, and the stargears can mesh with the ring gear 806. The sun gear 802 can be coupledto a low-speed shaft 808, which in turn is coupled to the low-pressureturbine of Engine 2. The carrier 804 can be fixed from rotation by asupport member 810. The ring gear 806 can be coupled to a fan shaft 812.

Based on the configuration of the gearbox 800 and the calculated oilflow rate of 4-13 gallons per minute, which is based on the gearboxefficiency rating, the gearbox 800 can comprise a radius R₂. The size ofthe gearbox, including the radius R₂, can be configured such that theoil flow rate at the inlet of the gearbox 800 at a max takeoff conditionis 7-13 gallons per minute (e.g., 10.1 gpm). In some embodiments, theradius R₂ of the gearbox 800 can be about 18-23 inches. In otherembodiments, the radius R₂ of the gearbox 700 can be smaller than 18inches or larger than 23 inches.

As another example, Engine 3 (FIG. 8) comprises an unducted fan and canbe configured similar to the engine 200. Engine 3 comprises a fandiameter of 142.8 inches and a net thrust of 12,500 pounds force at amax takeoff condition. Engine 3 further comprises a 3-7 stagelow-pressure turbine. The desired gear ratio for the gearbox of Engine 3is about 7:1. Based on this information, oil flow rate Q of the gearboxof Engine 3 should be about 3-9 gallons per minute or about 5-9 gallonsper minute (e.g., 6 gpm) at a max takeoff condition.

FIG. 11 schematically depicts a gearbox 900 that can be used, forexample, with Engine 3. The gearbox 800 comprises a single-stage starconfiguration. The gearbox 900 includes a sun gear 902, a carrier 904housing a plurality of star gears (e.g., 3-5 star gears), and a ringgear 906. The sun gear 902 can mesh with the star gears, and the stargears can mesh with the ring gear 906. The sun gear 902 can be coupledto a low-speed shaft 908, which in turn is coupled to the low-pressureturbine of Engine 3. The carrier 904 can be fixed from rotation by asupport member 910. The ring gear 906 can be coupled to a fan shaft 912.

Based on the configuration of the gearbox 900 and the calculated oilflow rate of 5-9 gallons per minute, which is based on the gearboxefficiency rating, the gearbox 900 can comprise a radius R₃. The size ofthe gearbox, including the radius R₃, can be configured such that theoil flow rate at the inlet of the gearbox 900 at a max takeoff conditionis 3-9 gallons per minute (e.g., 6 gpm). In some embodiments, the radiusR₃ of the gearbox 900 can be about 10-13 inches. In other embodiments,the radius R₃ of the gearbox 900 can be smaller than 10 inches or largerthan 13 inches.

Engine 4 comprises an unducted fan and can be configured similar to theengine 200. Engine 4 comprises a fan diameter of 188.4 inches and a netthrust of 25,000 pounds force at a max takeoff condition. Engine 4further comprises a counter-rotating low-pressure turbine (e.g., similarto the counter-rotating turbine 500 or the counter-rotating turbine600). The desired gear ratio for the gearbox of Engine 4 is about 7:1.Based on this information, oil flow rate Q of the gearbox of Engine 4should be about 4-13 gallons per minute or about 7-13 gallons per minute(e.g., 8.1 gpm) at a max takeoff condition.

FIG. 12 schematically depicts a gearbox 1000 that can be used, forexample, with Engine 4. The gearbox 1000 comprises a two-stageconfiguration in which the first stage is a star configuration and thesecond stage is a planet configuration.

The first stage of the gearbox 1000 includes a first-stage sun gear1002, a first-stage star carrier 1004 comprising a plurality offirst-stage star gears (e.g., 3-5 star gears), and a first-stage ringgear 1006. The first-stage sun gear 1002 can mesh with the first-stagestar gears, and the first-stage star gears can mesh with the first-stagering gear 1006. The first-stage sun gear 1002 can be coupled to ahigher-speed shaft 1008 of the low spool, which in turn is coupled tothe inner blades of the low-pressure turbine of Engine 4. Thefirst-stage star carrier 1004 can be fixed from rotation by a supportmember 1010.

The second stage of the gearbox 1000 includes a second-stage sun gear1012, a second-stage planet carrier 1014 comprising a plurality ofsecond-stage planet gears (e.g., 3-5 planet gears), and a second-stagering gear 1016. The second-stage sun gear 1012 can mesh with thesecond-stage planet gears. The second-stage planet carrier 1014 can becoupled to the first-stage ring gear 1006. The second-stage sun gear1012 can be coupled to a lower-speed shaft 1018 of the low spool, whichin turn is coupled to the outer blades of the low-pressure turbine ofEngine 4. The second-stage planet carrier 1014 can be coupled to thefirst-stage ring gear 1006. The second-stage planet carrier 1014 canalso be coupled to a fan shaft 1020. The second-stage ring gear 1016 canbe fixed from rotation by a support member 1022.

In some embodiments, each stage of the gearbox 1000 can comprise threestar/planet gears. In other embodiments, the gearbox 1000 can comprisefewer or more than three star/planet gears in each stage. In someembodiments, the first-stage carrier can comprise a different number ofstar gears than the second-stage carrier has planet gears. For example,the first-carrier can comprise five star gears, and the second-stagecarrier can comprise three planet gears, or vice versa.

Since the first stage of the gearbox 1000 is coupled to the higher-speedshaft 1008 of the low spool and the second stage of the gearbox 1000 iscoupled to the lower-speed shaft 1018 of the low spool, the gear ratioof the first stage of the gearbox 1000 can be greater than the gearratio of the second stage of the gearbox. For example, in certainembodiments, the first stage of the gearbox can comprise a gear ratio of4.1-14, and the second stage of the gearbox can comprise a gear ratiothat is less than the gear ratio of the first stage of the gearbox. Inparticular embodiments, the first stage of the gearbox can comprise agear ratio of 7, and the second stage of the gearbox can comprise a gearratio of 6.

In some embodiments, an engine comprising the gearbox 1000 can beconfigured such that the higher-speed shaft 1008 provides about 50% ofthe power to the gearbox 1000 and the lower-speed shaft 1018 providesabout 50% of the power to the gearbox 1000. In other embodiments, anengine comprising the gearbox 1000 can be configured such that thehigher-speed shaft 1008 provides about 60% of the power to the gearbox1000 and the lower-speed shaft 1018 provides about 40% of the power tothe gearbox 1000.

Based on the configuration of the gearbox 1000 and the calculated oilflow rate of 4-13 gallons per minute, which is based on the gearboxefficiency rating, the gearbox 1000 can comprise a radius R₄. The sizeof the gearbox, including the radius R₄, can be configured such that theoil flow rate at the inlet of the gearbox 1000 at a max takeoffcondition is 7-13 gallons per minute (e.g., 8.1 gpm). In someembodiments, the radius R₄ of the gearbox 1000 can be about 18-22inches. In other embodiments, the radius R₄ of the gearbox 700 can besmaller than 18 inches or larger than 22 inches.

Thus, as illustrated by the examples disclosed herein, a gearboxefficiency rating can characterize or define a specific engine and/orgearbox configuration. As such, turbomachine engines can be quickly andaccurately configured by utilizing the gearbox efficiency rating and/orits related parameters. In this manner, the gearbox efficiency ratingdisclosed herein provides one or more significant advantages over knownturbomachine engines and/or known methods of developing turbomachineengines.

FIG. 13 depicts a gearbox 1100 that can be used, for example, with theengines disclosed herein (e.g., the engines 100, 200, 400). The gearbox1100 is configured as a compound star gearbox. The gearbox 1100comprises a sun gear 1102 and a star carrier 1104, which includes aplurality of compound star gears having one or more first portions 1106and one or more second portions 1108. The gearbox 1100 further comprisesa ring gear 1110. The sun gear 1102 can also mesh with the firstportions 1106 of the star gears. The star carrier can be fixed fromrotation via a support member 1114. The second portions 1108 of the stargears can mesh with the ring gear 1110. The sun gear 1102 can be coupledto a low-pressure turbine via the turbine shaft 1112. The ring gear 1110can be coupled to a fan shaft 1116.

The gear assemblies shown and described herein can be used with anysuitable engine. For example, although FIG. 4 shows an optional ductedfan and optional fan duct (similar to that shown in FIG. 2), it shouldbe understood that such gear assemblies can be used with other ductedturbofan engines (e.g., the engine 300) and/or other open rotor enginesthat do not have one or more of such structures.

Embodiments of the gear assemblies depicted and described herein mayprovide for gear ratios and arrangements that fit within the L/D_(core)constraints of the disclosed engines. In certain embodiments, the gearassemblies depicted and described in regard to FIGS. 9-13 allow for gearratios and arrangements providing for rotational speed of the fanassembly corresponding to one or more ranges of cruise altitude and/orcruise speed provided above.

Various embodiments of the gear assembly provided herein may allow forgear ratios of up to 14:1. Still various embodiments of the gearassemblies provided herein may allow for gear ratios of at least 4.1:1or 4.5:1. Still yet various embodiments of the gear assemblies providedherein allow for gear ratios of 6:1 to 12:1. FIG. 8 also provides thegear ratio of several exemplary engines. It should be appreciated thatembodiments of the gear assemblies provided herein may allow for largegear ratios and within constraints such as, but not limited to, lengthof the engine, maximum diameter (D_(core)) of the engine 100, cruisealtitude of up to 65,000 ft, and/or operating cruise speed of up to Mach0.85, or combinations thereof.

Various exemplary gear assemblies are shown and described herein. Thesegear assemblies may be utilized with any of the exemplary engines and/orany other suitable engine for which such gear assemblies may bedesirable. In such a manner, it will be appreciated that the gearassemblies disclosed herein may generally be operable with an enginehaving a rotating element with a plurality of rotor blades and aturbomachine having a turbine and a shaft rotatable with the turbine.With such an engine, the rotating element (e.g., fan assembly 104) maybe driven by the shaft (e.g., low-speed shaft 146) of the turbomachinethrough the gear assembly.

Although the exemplary gear assemblies shown are mounted at a forwardlocation (e.g., forward from the combustor and/or the low-pressurecompressor), in other embodiments, the gear assemblies described hereincan be mounted at a aft location (e.g., aft of the combustor and/or thelow-pressure turbine).

Portions of a lubricant system 1200 are depicted schematically in FIG.14. The lubrication system 1200 can be a component of the turbomachineengines disclosed herein and/or can be coupled to the various gearboxesdisclosed herein. For example, FIG. 1 schematically illustrates thelubricant system coupled to the turbomachine engine 100 and the gearbox102. A series of lubricant conduits 1203 can interconnect multipleelements of the lubricant system 1200 and/or engine components, therebyproviding for provision or circulation of the lubricant throughout thelubricant system and any engine components coupled thereto (e.g., agearbox, bearing compartments, etc.).

It should be understood that the organization of the lubricant system1200 as shown is by way of example only to illustrate an exemplarysystem for a turbomachine engine for circulating lubricant for purposessuch as lubrication or heat transfer. Any organization for the lubricantsystem 1200 is contemplated, with or without the elements as shown,and/or including additional elements interconnected by any necessaryconduit system.

Referring again to FIG. 14, the lubricant system 1200 includes alubricant reservoir 1202 configured to store a coolant or lubricant,including organic or mineral oils, synthetic oils, or fuel, or mixturesor combinations thereof. A supply line 1204 and a scavenge line 1206 arefluidly coupled to the reservoir 1202 and collectively form a lubricantcircuit to which the reservoir 1202 and component 1210 (e.g., a gearbox)can be fluidly coupled. The component 1210 can be supplied withlubrication by way of a fluid coupling with the supply line 1204 and canreturn the supplied lubricant to the reservoir 1202 by fluidly couplingto the scavenge line 1206. More specifically, a component supply line1211 can be fluidly coupled between the supply line 1204 and thecomponent 1210. It is further contemplated that multiple types oflubricant can be provided in other lines not explicitly shown, but arenonetheless included in the lubricant system 1200.

Optionally, at least one heat exchanger 1205 can be included in thelubricant system 1200. The heat exchanger 1205 can include afuel/lubricant (fuel-to-lubricant) heat exchanger, an oil/lubricant heatexchanger, an air cooled oil cooler, and/or other means for exchangingheat. For example, a fuel/lubricant heat exchanger can be used to heator cool engine fuel with lubricant passing through the heat exchanger.In another example, a lubricant/oil heat exchanger can be used to heator cool additional lubricants passing within the turbomachine engine,fluidly separate from the lubricant passing along the lubricant system1200. Such a lubricant/oil heat exchanger can also include aservo/lubricant heat exchanger. Optionally, a second heat exchanger (notshown) can be provided along the exterior of the core engine, downstreamof the outlet guide vane assembly. The second heat exchanger can be anair/lubricant heat exchanger, for example, adapted to convectively coollubricant in the lubricant system 1200 utilizing the airflow passingthrough an outlet guide vane assembly of the turbomachine engine.

A pump 1208 can be provided in the lubricant system 1200 to aid inrecirculating lubricant from the reservoir 1202 to the component 1210via the supply line 1204. For example, the pump 1208 can be driven by arotating component of the turbine engine 10, such as a high-pressureshaft or a low-pressure shaft of a turbomachine engine.

Lubricant can be recovered from the component 1210 by way of thescavenge line 1206 and returned to the reservoir 1202. In theillustrated example, the pump 1208 is illustrated along the supply line1204 downstream of the reservoir 1202. The pump 1208 can be located inany suitable position within the lubricant system 1200, including alongthe scavenge line 1206 upstream of the reservoir 1202. In addition,while not shown, multiple pumps can be provided in the lubricant system1200.

In some embodiments, a bypass line 1212 can be fluidly coupled to thesupply line 1204 and scavenge line 1206 in a manner that bypasses thecomponent 1210. In such embodiments, a bypass valve 1215 is fluidlycoupled to the supply line 1204, component supply line 1211, and bypassline 1212. The bypass valve 1215 is configured to control a flow oflubricant through at least one of the component supply line 1211 or thebypass line 1212. The bypass valve 1215 can include any suitable valveincluding, but not limited to, a differential thermal valve, rotaryvalve, flow control valve, and/or pressure safety valve. In someembodiments, a plurality of bypass valves can be provided.

During operation, a supply flow 1220 can move from the reservoir 1202,through the supply line 1204, and to the bypass valve 1215. A componentinput flow 1222 can move from the bypass valve 1215 through thecomponent supply line 1211 to an inlet of the component 1210. A scavengeflow 1224 can move lubricant from an outlet of the component 1210through the scavenge line 1206 and back to the reservoir 1202.Optionally, a bypass flow 1226 can move from the bypass valve 1215through the bypass line 1212 and to the scavenge line 1206. The bypassflow 1226 can mix with the scavenge flow 1224 and define a return flow1228 moving toward the lubricant reservoir 1202.

In one example where no bypass flow exists, it is contemplated that thesupply flow 1220 can be the same as the component input flow 1222 andthat the scavenge flow 1224 can be the same as the return flow 1228. Inanother example where the bypass flow 1226 has a nonzero flow rate, thesupply flow 1220 can be divided at the bypass valve 1215 into thecomponent input flow 1222 and bypass flow 1226. It will also beunderstood that additional components, valves, sensors, or conduit linescan be provided in the lubricant system 1200, and that the example shownin FIG. 14 is simplified with a single component 1210 for purposes ofillustration.

The lubricant system 1200 can further include at least one sensingposition at which at least one lubricant parameter can be sensed ordetected. The at least one lubricant parameter can include, but is notlimited to, a flow rate, a temperature, a pressure, a viscosity, achemical composition of the lubricant, or the like. In the illustratedexample, a first sensing position 1216 is located in the supply line1204 upstream of the component 1210, and a second sensing position 1218is located in the scavenge line 1206 downstream of the component 1210.

In one example, the bypass valve 1215 can be in the form of adifferential thermal valve configured to sense or detect at least onelubricant parameter in the form of a temperature of the lubricant. Insuch a case, the fluid coupling of the bypass valve 1215 to the firstand second sensing positions 1216, 1218 can provide for bypass valve1215 sensing or detecting the lubricant temperature at the sensingpositions 1216, 18 as lubricant flows to or from the bypass valve 1215.The bypass valve 1215 can be configured to control the component inputflow 1222 or the bypass flow 1226 based on the sensed or detectedtemperature.

It is contemplated that the bypass valve 1215, supply line 1204, andbypass line 1212 can at least partially define a closed loop controlsystem for the component 1210. As used herein, a “closed loop controlsystem” will refer to a system having mechanical or electroniccomponents that can automatically regulate, adjust, modify, or control asystem variable without manual input or other human interaction. Suchclosed loop control systems can include sensing components to sense ordetect parameters related to the desired variable to be controlled, andthe sensed or detected parameters can be utilized as feedback in a“closed loop” manner to change the system variable and alter the sensedor detected parameters back toward a target state. In the example of thelubricant system 1200, the bypass valve 1215 (e.g. mechanical orelectrical component) can sense a parameter, such as the lubricantparameter (e.g. temperature), and automatically adjust a systemvariable, e.g., flow rate to either or both of the bypass line 1212 orcomponent 1210, without need of additional or manual input. In oneexample, the bypass valve can be automatically adjustable orself-adjustable such as a thermal differential bypass valve. In anotherexample, the bypass valve can be operated or actuated via a separatecontroller. It will be understood that a closed loop control system asdescribed herein can incorporate such a self-adjustable bypass valve ora controllable bypass valve.

Turning to FIG. 15, a portion of the lubricant system 1200 isillustrated supplying lubricant to a particular component 1210 in theform of a gearbox 1250 within a turbomachine engine. The gearbox can beany of the gearboxes disclosed herein. The gearbox 1250 can include aninput shaft 1252, an output shaft 1254, and a gear assembly 1255. In oneexample, the gear assembly 1255 can be in the form of an epicyclic gearassembly as known in the art having a ring gear, sun gear, and at leastone planet/star gear. An outer housing 1256 can at least partiallysurround the gear assembly 1255 and form a structural support for thegears and bearings therein. Either or both of the input and outputshafts 1252, 1254 can be coupled to the turbomachine engine. In oneexample, the input and output shafts 1252, 1254 can be utilized todecouple the speed of the low-pressure turbine from the low-pressurecompressor and/or the fan, which can, for example, improve engineefficiency.

The supply line 1204 can be fluidly coupled to the gearbox 1250, such asto the gear assembly 1255, to supply lubricant to gears or bearings tothe gearbox 1250 during operation. The scavenge line 1206 can be fluidlycoupled to the gearbox 1250, such as to the gear assembly 1255 or outerhousing 1256, to collect lubricant. The bypass line 1212 can be fluidlycoupled to the bypass valve 1215, supply line 1204, and scavenge line1206 as shown. A return line 1214 can also be fluidly coupled to thebypass valve 1215, such as for directing the return flow 1228 to thelubricant reservoir 1202 for recirculation. While not shown in FIG. 15for brevity, the lubricant reservoir 1202, the heat exchanger 1205,and/or the pump 1208 (FIG. 14) can also be fluidly coupled to thegearbox 1250. In this manner, the supply line 1204, bypass line 1212,scavenge line 1206, and return line 1214 can at least partially define arecirculation line 1230 for the lubricant system 1200.

The supply flow 1220 divides at the bypass line into the component inputflow 1222 and the bypass flow 1226. In the example shown, the bypassvalve 1215 is in the form of a differential thermal valve that isfluidly coupled to the first and second sensing positions 1216, 1218.

Lubricant flowing proximate the first and second sensing positions 1216,1218 provides the respective first and second outputs 1241, 1242indicative of the temperature of the lubricant at those sensingpositions 1216, 1218. It will be understood that the supply line 1204 isthermally coupled to the bypass line 1212 and bypass valve 1215 suchthat the temperature of the fluid in the supply line 1204 proximate thefirst sensing position 1216 is approximately the same as fluid in thebypass line 1212 adjacent the bypass valve 1215. Two values being“approximately the same” as used herein will refer to the two values notdiffering by more than a predetermined amount, such as by more than 20%,or by more than 5 degrees, in some examples. In this manner, the bypassvalve 1215 can sense the lubricant temperature in the supply line 1204and scavenge line 1206 via the first and second outputs 1241, 1242. Itcan be appreciated that the bypass line 1212 can form a sensing line forthe valve 1215 to sense the lubricant parameter, such as temperature, atthe first sensing position 1216.

During operation of the turbomachine engine, the lubricant temperaturecan increase within the gearbox 1250, such as due to heat generation ofthe gearbox 1250, and throughout the lubricant system 1200. In oneexample, if a lubricant temperature exceeds a predetermined thresholdtemperature at either sensing position 1216, 1218, the bypass valve 1215can automatically increase the component input flow 1222, e.g. from thesupply line 1204 to the gearbox 1250, by decreasing the bypass flow1226. Such a predetermined threshold temperature can be any suitableoperating temperature for the gearbox 1250, such as about 300° F. insome examples. Increasing the component input flow 1222 can provide forcooling of the gearbox 1250, thereby reducing the lubricant temperaturesensed in the various lines 1204, 1206, 1212, 1214 as lubricantrecirculates through the lubricant system 1200.

In another example, if a temperature difference between the sensingpositions 1216, 1218 exceeds a predetermined threshold temperaturedifference, the bypass valve can automatically increase the componentinput flow 1222 by decreasing the bypass flow 1226. Such a predeterminedthreshold temperature difference can be any suitable operatingtemperature for the gearbox 1250, such as about 70° F., or differing bymore than 30%, in some examples. In yet another example, if atemperature difference between the sensing positions 1216, 1218 is belowthe predetermined threshold temperature difference, the bypass valve canautomatically decrease the component input flow 1222 or increase thebypass flow 1226. In this manner the lubricant system 1200 can providefor the gearbox to operate with a constant temperature differencebetween the supply and scavenge lines 1204, 1206.

This written description uses examples to disclose the technology,including the best mode, and also to enable any person skilled in theart to practice the disclosed technology, including making and using anydevices or systems and performing any incorporated methods. Thepatentable scope of the disclosed technology is defined by the claims,and may include other examples that occur to those skilled in the art.Such other examples are intended to be within the scope of the claims ifthey include structural elements that do not differ from the literallanguage of the claims, or if they include equivalent structuralelements with insubstantial differences from the literal languages ofthe claims.

Further aspects of the disclosure are provided by the subject matter ofthe following clauses:

1. A turbomachine engine comprising a fan assembly including a pluralityof fan blades, a vane assembly including a plurality of vanes, a coreengine including one or more compressor sections and one or more turbinesections, a gearbox including an input and an output, and a gearboxefficiency rating of 0.10-1.8. The input is coupled to the one or moreturbine sections of the core engine and comprises a first rotationalspeed, the output is coupled to the fan assembly and has a secondrotational speed, and a gear ratio of the first rotational speed to thesecond rotational speed is within a range of 4.1-14.0. The gearboxefficiency rating equals Q(D{circumflex over ( )}1.56/T){circumflex over( )}1.53, where Q is a gearbox oil flow rate at an inlet of the gearboxmeasured in gallons per minute at a max takeoff condition, D is adiameter of the fan blades measured in inches, and T is a net thrust ofthe turbomachine engine measured in pounds force at the max takeoffcondition.

2. The turbomachine engine of any clause herein, wherein the gearboxefficiency rating is 0.10-1.01.

3. The turbomachine engine of any clause herein, wherein the gearboxefficiency rating is 0.19-1.8

4. The turbomachine engine of any clause herein, wherein the gear ratiois within a range of 4.5-12.0.

5. The turbomachine engine of any clause herein, wherein the gear ratiois within a range of 6.0-11.0.

6. The turbomachine engine of any clause herein, wherein Q is within arange of 5-55 gallons per minute.

7. The turbomachine engine of any clause herein, wherein Q is within arange of 6-36 gallons per minute.

8. The turbomachine engine of any clause herein, wherein D is 120-216inches.

9. The turbomachine engine of any clause herein, wherein D is 120-192inches.

10. The turbomachine engine of any clause herein, wherein T is within arange of 10,000-100,000 pounds force.

11. The turbomachine engine of any clause herein, wherein T is within arange of 12,000-30,000 pounds force.

12. The turbomachine engine of any clause herein, wherein the gearbox isan epicyclic gearbox comprising a sun gear, a plurality of planet gears,and a ring gear, wherein the sun gear is the input, and wherein the ringgear is the output.

13. The turbomachine engine of any clause herein, wherein the gearbox isan epicyclic gearbox comprising a sun gear, a plurality of planet gears,and a ring gear, wherein the sun gear is the input, wherein the planetgears are coupled to a planet carrier, and wherein the planet carrier isthe output.

14. The turbomachine engine of any clause herein, wherein the gearbox isa multi-stage gearbox.

15. The turbomachine engine of any clause herein, wherein the gearbox isa two-stage gearbox.

16. The turbomachine engine of any clause herein, wherein the gearbox isa compound gearbox.

17. A turbomachine engine comprising a fan assembly including aplurality of fan blades, a vane assembly including a plurality of vanes,a core engine including a low-pressure compressor, a high-pressurecompressor, a combustor, a high-pressure turbine, and a low-pressureturbine, a gearbox including an input and an output, and a gearboxefficiency rating of 0.12-1.8. The input is coupled to the low-pressureturbine and comprises a first rotational speed, the output is coupled tothe fan assembly and has a second rotational speed, and a gear ratio ofthe first rotational speed to the second rotational speed is within arange of 4.5-14.0. The gearbox efficiency rating equals Q(D{circumflexover ( )}1.56/T){circumflex over ( )}1.53, where Q is a gearbox oil flowrate at an inlet of the gearbox measured in gallons per minute at a maxtakeoff condition, D is a diameter of the fan blades measured in inches,and T is a net thrust of the turbomachine engine measured in poundsforce at the max takeoff condition.

18. The turbomachine engine of any clause herein, wherein Q is within arange of 5-55 gallons per minute.

19. The turbomachine engine of any clause herein, wherein D is 120-216inches.

20. The turbomachine engine of any clause herein, wherein T is within arange of 10,000-100,000 pounds force.

21. The turbomachine engine of any clause herein, wherein the gearbox isan epicyclic gearbox comprising a star gear configuration.

22. The turbomachine engine of any clause herein, wherein the gearbox isan epicyclic gearbox comprising a planet gear configuration.

23. The turbomachine engine of any clause herein, wherein the fanassembly comprises 8-20 fan blades.

24. The turbomachine engine of any clause herein, wherein thelow-pressure compressor comprises 1-8 stages.

25. The turbomachine engine of any clause herein, wherein thehigh-pressure compressor comprises 8-15 stages.

26. The turbomachine engine of any clause herein, wherein thehigh-pressure turbine comprises 1-2 stages.

27. The turbomachine engine of any clause herein, wherein thelow-pressure turbine comprises 3-7 stages.

28. The turbomachine engine of any clause herein, wherein thelow-pressure turbine is a counter-rotating low-pressure turbinecomprising inner blade stages and outer blade stages, wherein the innerblade stages extend radially outwardly from an inner shaft, and whereinthe outer blade stages extend radially inwardly from an outer drum.

29. The turbomachine engine of any clause herein, wherein thecounter-rotating low-pressure turbine comprises four inner blade stagesand three outer blade stages.

30. The turbomachine engine of any clause herein, wherein thecounter-rotating low-pressure turbine comprises three inner blade stagesand three outer blade stages.

31. A turbomachine engine comprising a fan assembly including aplurality of fan blades, a vane assembly including a plurality of vanes,a core engine including a low-pressure compressor, a high-pressurecompressor, a combustor, a high-pressure turbine, and a low-pressureturbine. The turbomachine engine also includes a gearbox having a gearratio within a range of 6.0-12.0, and a gearbox efficiency rating of0.18-1.41.

32. The turbomachine engine of any clause herein, wherein a gearbox oilflow rate at an inlet of the gearbox is within a range of 5-55 gallonsper minute at a max takeoff condition.

33. The turbomachine engine of any clause herein, wherein a diameter ofthe fan blades is 120-216 inches.

34. The turbomachine engine of any clause herein, wherein a net thrustof the turbomachine engine is within a range of 10,000-100,000 poundsforce at a max takeoff condition.

35. A turbomachine engine comprises an unducted fan assembly, a coreengine, a vane assembly, a gearbox, and a gearbox efficiency rating. Theunducted fan assembly includes a single row of fan blades. The coreengine including one or more compressor sections and one or more turbinesections. The vane assembly includes a single row of vanes. The vanesare disposed aft of the fan blades and comprise fixed end portions andfree end portions. The fixed end portions are coupled to the coreengine, and the free end portions are spaced radially outwardly from thecore engine. The gearbox includes an input and an output. The input iscoupled to the one or more turbine sections of the core engine andcomprises a first rotational speed, the output is coupled to theunducted fan assembly and has a second rotational speed, and a gearratio of the first rotational speed to the second rotational speed iswithin a range of 4.1-14.0. The gearbox efficiency rating is 0.10-1.8.The gearbox efficiency rating equals

${Q\left( \frac{D^{1.56}}{T} \right)}^{1.53},$where Q is a gearbox oil flow rate at an inlet of the gearbox measuredin gallons per minute at a max takeoff condition, D is a diameter of thefan blades measured in inches, and T is a net thrust of the turbomachineengine measured in pounds force at the max takeoff condition.

36. The turbomachine engine of any clause herein, further comprising apitch change mechanism coupled to the unducted fan assembly.

37. The turbomachine engine of any clause herein, further comprising apitch change mechanism coupled to the vane assembly.

38. The turbomachine engine of any clause herein, wherein the gearboxefficiency rating is 0.25-1.15.

39. The turbomachine engine of any clause herein, wherein the gear ratiois within a range of 4.5-12.0.

40. The turbomachine engine of any clause herein, wherein the gear ratiois within a range of 6.0-11.0.

41. The turbomachine engine of any clause herein, wherein Q is within arange of 5-55 gallons per minute.

42. The turbomachine engine of any clause herein, wherein D is 120-216inches.

43. The turbomachine engine of any clause herein, wherein T is within arange of 10,000-100,000 pounds force.

44. The turbomachine engine of any clause herein, wherein the gearbox isan epicyclic gearbox comprising a sun gear, a plurality of planet gears,and a ring gear, wherein the sun gear is the input, and wherein the ringgear is the output.

45. The turbomachine engine of any clause herein, wherein the gearbox isan epicyclic gearbox comprising a sun gear, a plurality of planet gears,and a ring gear, wherein the sun gear is the input, wherein the planetgears are coupled to a planet carrier, and wherein the planet carrier isthe output.

46. The turbomachine engine of any clause herein, wherein the gearbox isa multi-stage gearbox.

47. The turbomachine engine of any clause herein, wherein the gearbox isa compound gearbox.

48. A turbomachine engine comprises an unducted fan assembly, anunducted vane assembly, a ducted fan assembly, a core engine, a gearbox,and a gearbox efficiency rating. The unducted fan assembly includes aplurality of first fan blades. The unducted vane assembly including aplurality of vanes, and the vanes are positioned aft of the first fanblades. The ducted fan assembly includes a plurality of second fanblades, and the ducted fan assembly is positioned aft of the unductedfan assembly and radially inwardly from the unducted vane assembly. Thecore engine including a low-pressure compressor, a high-pressurecompressor, a combustor, a high-pressure turbine, and a low-pressureturbine. The gearbox is coupled to the low-pressure turbine and theunducted fan assembly. The gearbox comprises a gear ratio of 4.1-14.0and is configured such that the unducted fan assembly rotates slowerthan low-pressure turbine. The gearbox efficiency rating is 0.10-1.8.The gearbox efficiency rating equals

${Q\left( \frac{D^{1.56}}{T} \right)}^{1.53},$where Q is a gearbox oil flow rate at an inlet of the gearbox measuredin gallons per minute at a max takeoff condition, D is a diameter of thefan blades measured in inches, and T is a net thrust of the turbomachineengine measured in pounds force at the max takeoff condition.

49. The turbomachine engine of any clause herein, wherein the gearboxefficiency rating is 0.10-1.01.

50. The turbomachine engine of any clause herein, wherein the gearboxefficiency rating is 0.19-1.8.

51. The turbomachine engine of any clause herein, wherein Q is within arange of 5-40 gallons per minute.

52. The turbomachine engine of any clause herein, wherein D is 140-192inches.

53. The turbomachine engine of any clause herein, wherein T is within arange of 10,000-40,000 pounds force.

54. The turbomachine engine of any clause herein, wherein the gearbox isan epicyclic gearbox comprising a star gear configuration.

55. The turbomachine engine of any clause herein, wherein the gearbox isan epicyclic gearbox comprising a planet gear configuration.

56. The turbomachine engine of any clause herein, wherein the unductedfan assembly comprises 8-14 fan blades.

57. The turbomachine engine of any clause herein, wherein thelow-pressure compressor comprises 1-2 stages.

58. The turbomachine engine of any clause herein, wherein thehigh-pressure compressor comprises 10-11 stages.

59. The turbomachine engine of any clause herein, wherein thehigh-pressure turbine comprises two stages.

60. The turbomachine engine of any clause herein, wherein thelow-pressure turbine comprises 6-7 stages.

61. The turbomachine engine of any clause herein, wherein thelow-pressure turbine is a counter-rotating low-pressure turbinecomprising inner blade stages and outer blade stages, wherein the innerblade stages extend radially outwardly from an inner shaft, and whereinthe outer blade stages extend radially inwardly from an outer drum.

62. The turbofan engine of any clause herein, wherein the unducted fanassembly is configured to direct a first portion of airflow to theunducted vane assembly and a second portion of airflow into an inletduct and to the ducted fan assembly, and wherein the ducted fan assemblyis configured to direct the second portion of airflow to a fan duct andto a core duct.

63. A turbomachine engine comprises an open rotor fan assembly, a coreengine, a vane assembly, a gearbox, a gearbox efficiency rating. Theopen rotor fan assembly including a plurality of fan blades. The coreengine including a low-pressure compressor, a high-pressure compressor,a combustor, a high-pressure turbine, and a low-pressure turbine. Thevane assembly including a plurality of vanes extending radiallyoutwardly from the core engine in a cantilever manner. The gearbox iscoupled to the low-pressure turbine and the open rotor fan assembly. Thegearbox comprises a gear ratio of 6.0-12.0 and is configured such that afirst rotational speed of the open rotor fan assembly is less than asecond rotational speed of the low-pressure turbine. The gearboxefficiency rating is 0.18-1.41.

64. The turbomachine engine of any clause herein, wherein a gearbox oilflow rate at an inlet of the gearbox is within a range of 6-36 gallonsper minute at a max takeoff condition, wherein a diameter of the fanblades is 140-192 inches, and wherein a net thrust of the turbomachineengine is within a range of 12,000-30,000 pounds force at a max takeoffcondition.

65. A turbomachine engine comprises a fan case, a fan assembly, a pitchchange mechanism, a core engine, a vane assembly, a gearbox, and agearbox efficiency rating. The fan assembly is disposed radially withinthe fan case and comprises a plurality of fan blades. The pitch changemechanism is coupled to the fan assembly and is configured to adjust apitch of the fan blades. The core engine including a low-pressureturbine. The vane assembly includes a plurality of vanes. The vanes aredisposed aft of the fan blades and are coupled to the core engine andthe fan case. The gearbox is coupled to the low-pressure turbine and thefan assembly. The gearbox is configured such that a ratio of a firstrotational speed of the low-pressure turbine to a second rotationalspeed of the fan assembly is within a range of 4.1-14.0. The gearboxefficiency rating is 0.10-1.8. The gearbox efficiency rating equals

${Q\left( \frac{D^{1.56}}{T} \right)}^{1.53},$where Q is a gearbox oil flow rate at an inlet of the gearbox measuredin gallons per minute at a max takeoff condition, D is a diameter of thefan blades measured in inches, and T is a net thrust of the turbomachineengine measured in pounds force at the max takeoff condition.

66. The turbomachine engine of any clause herein, wherein the pitchchange mechanism is a first pitch change mechanism, and wherein theturbomachine engine further comprises a second pitch change mechanismcoupled to the vane assembly and configured to adjust a pitch of thevanes.

67. The turbomachine engine of any clause herein, wherein the gearboxefficiency rating is 0.25-1.5.

68. The turbomachine engine of any clause herein, wherein the ratio ofthe first rotational speed of the low-pressure turbine to the secondrotational speed of the fan assembly is within a range of 4.5-12.0.

69. The turbomachine engine of any clause herein, wherein the ratio ofthe first rotational speed of the low-pressure turbine to the secondrotational speed of the fan assembly is within a range of 6.0-11.0.

70. The turbomachine engine of any clause herein, wherein Q is within arange of 5-55 gallons per minute.

71. The turbomachine engine of any clause herein, wherein D is 120-216inches.

72. The turbomachine engine of any clause herein, wherein T is within arange of 10,000-100,000 pounds force.

73. The turbomachine engine of any clause herein, wherein the gearbox isan epicyclic gearbox comprising a sun gear, a plurality of planet gears,and a ring gear, wherein the sun gear is coupled to the low-pressureturbine, and wherein the ring gear is coupled to the fan assembly.

74. The turbomachine engine of any clause herein, wherein the gearbox isan epicyclic gearbox comprising a sun gear, a plurality of planet gears,and a ring gear, wherein the sun gear is coupled to the low-pressureturbine, wherein the planet gears are coupled to a planet carrier, andwherein the planet carrier is coupled to the fan assembly.

75. The turbomachine engine of any clause herein, wherein the gearbox isa multi-stage gearbox.

76. The turbomachine engine of any clause herein, wherein the gearbox isa compound gearbox.

77. A turbomachine engine comprises a fan case, a fan assembly, a pitchchange mechanism, a vane assembly, a core engine, a gearbox, and agearbox efficiency rating. The fan assembly is housed within the fancase and comprising a plurality of fan blades. The pitch changemechanism is coupled to the fan assembly. The vane assembly is housedwithin the fan case and comprises a plurality of vanes. The core engineincludes a low-pressure compressor, a high-pressure compressor, acombustor, a high-pressure turbine, and a low-pressure turbine. Thegearbox is coupled to the low-pressure turbine and the fan assembly. Thegearbox comprises a gear ratio of 4.1-14.0 and is configured such thatthe fan assembly rotates slower than one or more stages of thelow-pressure turbine. The gearbox efficiency rating is 0.10-1.8. Thegearbox efficiency rating equals

${Q\left( \frac{D^{1.56}}{T} \right)}^{1.53},$where Q is a gearbox oil flow rate at an inlet of the gearbox measuredin gallons per minute at a max takeoff condition, D is a diameter of thefan blades measured in inches, and T is a net thrust of the turbomachineengine measured in pounds force at the max takeoff condition.

77. The turbomachine engine of any clause herein, wherein the gearboxefficiency rating is 0.10-1.01.

78. The turbomachine engine of any clause herein, wherein the gearboxefficiency rating is 0.19-1.8

79. The turbomachine engine of any clause herein, wherein Q is within arange of 6-36 gallons per minute.

80. The turbomachine engine of any clause herein, wherein D is 140-192inches.

81. The turbomachine engine of any clause herein, wherein T is within arange of 10,000-40,000 pounds force.

82. The turbomachine engine of any clause herein, wherein the gearbox isan epicyclic gearbox comprising a star gear configuration.

83. The turbomachine engine of any clause herein, wherein the gearbox isan epicyclic gearbox comprising a planet gear configuration.

84. The turbomachine engine of any clause herein, wherein the fanassembly comprises 8-14 fan blades.

85. The turbomachine engine of any clause herein, wherein thelow-pressure compressor comprises 1-2 stages.

86. The turbomachine engine of any clause herein, wherein thehigh-pressure compressor comprises 10-11 stages.

87. The turbomachine engine of any clause herein, wherein thehigh-pressure turbine comprises two stages.

88. The turbomachine engine of any clause herein, wherein thelow-pressure turbine comprises 6-7 stages.

89. The turbomachine engine of any clause herein, wherein thelow-pressure turbine is a counter-rotating low-pressure turbinecomprising inner blade stages and outer blade stages, wherein the innerblade stages extend radially outwardly from an inner shaft, wherein theouter blade stages extend radially inwardly from an outer drum, andwherein the gearbox is configured such that the fan assembly rotatesslower than the inner blade stages of the low-pressure turbine.

90. A turbomachine engine comprises a fan case, a fan assembly, a pitchchange mechanism, a core engine, a vane assembly, a gearbox, and agearbox efficiency rating. The fan assembly includes a plurality of fanblades. The pitch change mechanism is coupled to the fan assembly. Thecore engine includes a low-pressure compressor, a high-pressurecompressor, a combustor, a high-pressure turbine, and a low-pressureturbine. The vane assembly includes a plurality of vanes. The gearbox iscoupled to the low-pressure turbine and the fan assembly. The gearboxcomprises a gear ratio of 6.0-12.0 and is configured such that a firstrotational speed of the fan assembly is less than a second rotationalspeed of the low-pressure turbine. The gearbox efficiency rating of0.18-1.41.

91. The turbomachine engine of any clause herein, wherein a gearbox oilflow rate at an inlet of the gearbox is within a range of 5-40 gallonsper minute at a max takeoff condition, wherein a diameter of the fanblades is 140-192 inches, and wherein a net thrust of the turbomachineengine is within a range of 12,000-30,000 pounds force at a max takeoffcondition.

92. A turbomachine engine comprises a fan case, a fan assembly, a vaneassembly, a core engine, a gearbox, and a gearbox efficiency rating. Thefan assembly comprises a plurality of fan blades. The vane assemblyincludes a plurality of vanes, and the vanes are disposed aft of the fanblades. The core engine includes a counter-rotating low-pressureturbine. The gearbox is coupled to the counter-rotating low-pressureturbine and the fan assembly. The gearbox is configured such that aratio of a first rotational speed of the counter-rotating low-pressureturbine to a second rotational speed of the fan assembly is within arange of 4.1-14.0. The gearbox efficiency rating is 0.10-1.8. Thegearbox efficiency rating equals

${Q\left( \frac{D^{1.56}}{T} \right)}^{1.53},$where Q is a gearbox oil flow rate at an inlet of the gearbox measuredin gallons per minute at a max takeoff condition, D is a diameter of thefan blades measured in inches, and T is a net thrust of the turbomachineengine measured in pounds force at the max takeoff condition.

93. The turbomachine engine of any clause herein, wherein the gearboxefficiency rating is 0.25-1.15.

94. The turbomachine engine of any clause herein, wherein the ratio ofthe first rotational speed of the counter-rotating low-pressure turbineto the second rotational speed of the fan assembly is within a range of4.5-12.0.

95. The turbomachine engine of any clause herein, wherein the ratio ofthe first rotational speed of the counter-rotating low-pressure turbineto the second rotational speed of the fan assembly is within a range of6.0-11.0.

96. The turbomachine engine of any clause herein, wherein Q is within arange of 5-55 gallons per minute.

97. The turbomachine engine of any clause herein, wherein D is 120-216inches.

98. The turbomachine engine of any clause herein, wherein T is within arange of 10,000-100,000 pounds force.

99. The turbomachine engine of any clause herein, wherein thecounter-rotating low-pressure turbine includes an inner rotor and anouter drum, wherein the inner rotor comprises a plurality of inner bladestages, wherein the outer drum comprises a plurality of outer bladestages, and wherein the outer blade stages are disposed between adjacentinner blade stages.

100. The turbomachine engine of any clause herein, wherein thecounter-rotating low-pressure turbine comprises exactly three innerblade stages and exactly three outer blade stages.

101. The turbomachine engine of any clause herein, wherein thecounter-rotating low-pressure turbine comprises exactly four inner bladestages and exactly three outer blade stages.

102. The turbomachine engine of any clause herein, wherein the gearboxis an epicyclic gearbox comprising a sun gear, a plurality of planetgears, and a ring gear, wherein the sun gear is coupled to the innerrotor of the counter-rotating low-pressure turbine, and wherein the ringgear is coupled to outer drum of the counter-rotating low-pressureturbine and the fan assembly.

103. The turbomachine engine of any clause herein, wherein the gearboxis a multi-stage gearbox comprising a first stage and a second stage,wherein the first stage of the gearbox comprises a first-stage sun gear,a plurality of first-stage planet gears coupled to a first-stage planetcarrier, and a first-stage ring gear, wherein the second stage of thegearbox comprises a second-stage sun gear, a plurality of second-stageplanet gears coupled to a second-stage planet carrier, and asecond-stage ring gear, wherein the first-stage sun gear is coupled tothe inner rotor of the counter-rotating low-pressure turbine, andwherein second-stage sun gear is coupled to the outer drum of thecounter-rotating low-pressure turbine.

104. The turbomachine engine of any clause herein, wherein the firststage of the gearbox comprises a star gear configuration, and whereinthe second stage of the gearbox comprises a planet gear configuration.

105. The turbomachine engine of any clause herein, further comprising apitch change mechanism coupled to the fan assembly and configured toadjust a pitch of the fan blades.

106. A turbomachine engine comprises a fan case, a fan assembly, a vaneassembly, a core engine, a gearbox, and a gearbox efficiency rating. Thefan assembly is housed within the fan case and comprises a plurality offan blades. The vane assembly is housed within the fan case andcomprises a plurality of vanes. The core engine includes a low-pressurecompressor, a high-pressure compressor, a combustor, a high-pressureturbine, and a counter-rotating low-pressure turbine. The gearbox iscoupled to the counter-rotating low-pressure turbine and the fanassembly. The gearbox comprises a gear ratio of 4.1-14.0 and isconfigured such that the fan assembly rotates slower than one or morestages of the counter-rotating low-pressure turbine. The gearboxefficiency rating is 0.10-1.8. The gearbox efficiency rating equals

${Q\left( \frac{D^{1.56}}{T} \right)}^{1.53},$where Q is a gearbox oil flow rate at an inlet of the gearbox measuredin gallons per minute at a max takeoff condition, D is a diameter of thefan blades measured in inches, and T is a net thrust of the turbomachineengine measured in pounds force at the max takeoff condition.

107. The turbomachine engine of any clause herein, wherein the gearboxis an epicyclic gearbox comprising a star gear configuration.

108. The turbomachine engine of any clause herein, wherein the gearboxcomprises a first stage and a second stage, and wherein the fan assemblyrotates slower than all stages of the counter-rotating low-pressureturbine.

109. The turbomachine engine of any clause herein, wherein the firststage of the gearbox comprises a star gear configuration, and whereinthe second stage of the gearbox comprises a planet gear configuration.

110. The turbomachine engine of any clause herein, wherein the gearboxefficiency rating is 0.20-1.10.

111. The turbomachine engine of any clause herein, wherein the gearboxefficiency rating is 0.10-1.01.

112. The turbomachine engine of any clause herein, wherein the gearboxefficiency rating is 0.19-1.8.

113. The turbomachine engine of any clause herein, wherein Q is within arange of 6-36 gallons per minute.

114. The turbomachine engine of any clause herein, wherein D is 140-192inches.

115. The turbomachine engine of any clause herein, wherein T is within arange of 10,000-40,000 pounds force.

116. The turbomachine engine of any clause herein, wherein the fanassembly comprises 8-14 fan blades.

117. The turbomachine engine of any clause herein, wherein thelow-pressure compressor comprises 1-2 stages.

118. The turbomachine engine of any clause herein, wherein thehigh-pressure compressor comprises 10-11 stages.

119. The turbomachine engine of any clause herein, wherein thehigh-pressure turbine comprises two stages.

120. The turbomachine engine of any clause herein, wherein thecounter-rotating low-pressure turbine comprises 6-7 stages.

121. The turbomachine engine of any clause herein, wherein thecounter-rotating low-pressure turbine comprises inner blade stages andouter blade stages, wherein the inner blades stages are coupled to afirst rotatable shaft, and wherein the outer blades stages are coupledto a second rotatable shaft.

122. The turbomachine engine of any clause herein, wherein the gearboxis located forward from the combustor.

123. The turbomachine engine of any clause herein, wherein the gearboxis located aft of the combustor.

124. A turbomachine engine comprises a ducted fan assembly, a pitchchange mechanism, a core engine, a ducted vane assembly, a gearbox, anda gearbox efficiency rating. The ducted fan assembly includes aplurality of fan blades. The pitch change mechanism is coupled to theducted fan assembly. The core engine includes a low-pressure compressor,a high-pressure compressor, a combustor, a high-pressure turbine, and acounter-rotating low-pressure turbine. The ducted vane assembly includesa plurality of vanes. The gearbox is coupled to the counter-rotatinglow-pressure turbine and the ducted fan assembly. The gearbox comprisesa gear ratio of 6.0-12.0 and is configured such that a first rotationalspeed of the ducted fan assembly is less than a second rotational speedof one or more stages of the counter-rotating low-pressure turbine. Thegearbox efficiency rating is 0.18-1.41 at a max takeoff condition.

125. The turbomachine engine of any clause herein, wherein a gearbox oilflow rate at an inlet of the gearbox is within a range of 6-36 gallonsper minute at a max takeoff condition, wherein a diameter of the fanblades is 140-192 inches, and wherein a net thrust of the turbomachineengine is within a range of 12,000-30,000 pounds force at a max takeoffcondition.

126. A turbomachine engine comprises a fan case, a fan assembly, a vaneassembly, a core engine, a gearbox, and a gearbox efficiency rating. Thefan assembly is disposed radially within the fan case and comprises aplurality of fan blades. The core engine includes a low-pressureturbine. The vane assembly includes a plurality of vanes, and the vanesare disposed aft of the fan blades and are coupled to the core engineand the fan case. The gearbox is coupled to the low-pressure turbine andthe fan assembly, and the gearbox comprises a gear ratio within a rangeof 4.1-14.0. The gearbox efficiency rating is 0.10-1.8 at a max takeoffcondition.

127. The turbomachine engine of any clause herein, wherein the gearboxefficiency rating is 0.25-1.15 at the max takeoff condition.

128. The turbomachine engine of any clause herein, wherein the gearboxefficiency rating is 0.10-1.01 at the max takeoff condition.

129. The turbomachine engine of any clause herein, wherein the gearboxefficiency rating is 0.19-1.8 at the max takeoff condition.

130. The turbomachine engine of any clause herein, wherein the gearratio of the gearbox is within a range of 4.5-12.0.

131. The turbomachine engine of any clause herein, wherein the gearratio of the gearbox is within a range of 6.0-11.0.

132. The turbomachine engine of any clause herein, wherein a gearbox oilflow rate at an inlet of the gearbox is within a range of 5-55 gallonsper minute at the max takeoff condition.

133. The turbomachine engine of any clause herein, wherein a diameter ofthe fan blades is 72-216 inches.

134. The turbomachine engine of any clause herein, wherein a net thrustof the turbomachine engine is within a range of 10,000-100,000 poundsforce at the max takeoff condition.

135. The turbomachine engine of any clause herein, wherein the gearboxis an epicyclic gearbox comprising a sun gear, a plurality of planetgears, and a ring gear, wherein the sun gear is coupled to thelow-pressure turbine, and wherein the ring gear is coupled to the fanassembly.

136. The turbomachine engine of any clause herein, wherein the gearboxis an epicyclic gearbox comprising a sun gear, a plurality of planetgears, and a ring gear, wherein the sun gear is coupled to thelow-pressure turbine, wherein the planet gears are coupled to a planetcarrier, and wherein the planet carrier is coupled to the fan assembly.

137. The turbomachine engine of any clause herein, wherein the gearboxis a multi-stage gearbox.

138. The turbomachine engine of any clause herein, wherein the gearboxcomprises one or more compound gears, wherein each compound gearincludes a first portion having a first diameter and a second portionhaving a second diameter, the second diameter being less than the firstdiameter.

139. The turbomachine engine of any clause herein, further comprisingone or more pitch change mechanisms coupled to the fan assembly or thevane assembly.

140. A turbomachine engine comprises a fan case, a fan assembly, a vaneassembly, a core engine, a gearbox, and a gearbox efficiency rating. Thefan assembly is housed within the fan case and comprising a plurality offan blades. The vane assembly is housed within the fan case andcomprising a plurality of vanes. The core engine includes a low-pressurecompressor, a high-pressure compressor, a combustor, a high-pressureturbine, and a low-pressure turbine. The gearbox is coupled to thelow-pressure turbine and the fan assembly. The gearbox comprises a gearratio of 4.1-14.0 and is configured such that the fan assembly rotatesslower than one or more stages of the low-pressure turbine. The gearboxefficiency rating of 0.10-1.8 at a max takeoff condition.

141. The turbomachine engine of any clause herein, wherein the gearboxefficiency rating is 0.20-1.15 at the max takeoff condition.

142. The turbomachine engine of any clause herein, wherein a gearbox oilflow rate at an inlet of the gearbox is within a range of 6-36 gallonsper minute at the max takeoff condition.

143. The turbomachine engine of any clause herein, wherein the fanblades comprise a diameter within a range of 72-120 inches.

144. The turbomachine engine of any clause herein, wherein a net thrustof the turbomachine engine is within a range of 10,000-40,000 poundsforce at the max takeoff condition.

145. The turbomachine engine of any clause herein, wherein the gearboxis an epicyclic gearbox comprising a star gear configuration.

146. The turbomachine engine of any clause herein, wherein the gearboxis an epicyclic gearbox comprising a planet gear configuration.

147. The turbomachine engine of any clause herein, wherein the fanassembly comprises 8-20 fan blades.

148. The turbomachine engine of any clause herein, wherein thelow-pressure compressor comprises 3-8 stages.

149. The turbomachine engine of any clause herein, wherein thehigh-pressure compressor comprises 8-15 stages.

150. The turbomachine engine of any clause herein, wherein thehigh-pressure turbine comprises 1-2 stages.

151. The turbomachine engine of any clause herein, wherein thelow-pressure turbine comprises 3-6 stages.

152. The turbomachine engine of any clause herein, wherein thelow-pressure turbine is a counter-rotating low-pressure turbine.

153. A turbomachine engine comprises a fan case, a fan assembly, a vaneassembly, a core engine, an epicyclic gearbox, and a gearbox efficiencyrating. The fan assembly is housed within the fan case and comprises16-20 fan blades. The vane assembly is housed within the fan case andcomprises a plurality of vanes. The core engine includes a low-pressurecompressor, a high-pressure compressor, a combustor, a high-pressureturbine, and a low-pressure turbine. The low-pressure compressorcomprises 2-4 stages, the high-pressure compressor comprises 8-10stages, the high-pressure turbine comprises two stages, and thelow-pressure turbine comprises 3-4 stages. The epicyclic gearbox iscoupled to the low-pressure turbine and the fan assembly. The epicyclicgearbox comprises a gear ratio of 4.1-14.0 and is configured such thatthe fan assembly rotates slower than one or more stages of thelow-pressure turbine. The gearbox efficiency rating is 0.10-1.8 at a maxtakeoff condition.

154. The turbomachine engine of any clause herein, wherein the gearboxefficiency rating is 0.25-0.55 at the max takeoff condition.

155. The turbomachine engine of any clause herein, wherein the gearratio of the epicyclic gearbox is within a range of 6.0-12.0.

156. The turbomachine engine of any clause herein, wherein a gearbox oilflow rate at an inlet of the epicyclic gearbox is within a range of 5-40gallons per minute at the max takeoff condition.

157. The turbomachine engine of any clause herein, wherein a diameter ofthe fan blades is 72-120 inches.

158. The turbomachine engine of any clause herein, wherein a net thrustof the turbomachine engine is within a range of 10,000-40,000 poundsforce at the max takeoff condition.

159. The turbomachine engine of any clause herein, wherein the epicyclicgearbox comprises a sun gear, a plurality of planet gears, and a ringgear, wherein the sun gear is coupled to the low-pressure turbine, andwherein the ring gear is coupled to the fan assembly.

160. The turbomachine engine of any clause herein, wherein the epicyclicgearbox comprises a sun gear, a plurality of planet gears, and a ringgear, wherein the sun gear is coupled to the low-pressure turbine,wherein the planet gears are coupled to a planet carrier, and whereinthe planet carrier is coupled to the fan assembly.

161. The turbomachine engine of any clause herein, wherein the epicyclicgearbox is a multi-stage gearbox.

162. The turbomachine engine of any clause herein, further comprisingone or more pitch change mechanisms coupled to the fan assembly or thevane assembly.

163. The turbomachine engine of any clause herein, wherein the gearboxis a high power gearbox.

The invention claimed is:
 1. A turbomachine engine comprising: a fanassembly including a plurality of fan blades arranged in a single row; avane assembly including a plurality of vanes, wherein the vane assemblyis disposed aft of the fan assembly and configured to receive bypassairflow from the fan assembly; a core engine including one or morecompressor sections and one or more turbine sections; a gearboxincluding an input and an output, wherein the input is coupled to theone or more turbine sections of the core engine and comprises a firstrotational speed, wherein the output is coupled to the fan assembly andhas a second rotational speed, and wherein a gear ratio of the firstrotational speed to the second rotational speed is within a range of4.1-14.0; and a gearbox efficiency rating of 0.10-1.8, wherein thegearbox efficiency rating equals${Q\left( \frac{D^{1.56}}{T} \right)}^{1.53},$ wherein Q is a gearboxoil flow rate at an inlet of the gearbox measured in gallons per minuteat a max takeoff condition, wherein D is a diameter of the fan bladesmeasured in inches, and wherein T is a net thrust of the turbomachineengine measured in pounds force at the max takeoff condition.
 2. Theturbomachine engine of claim 1, wherein the gearbox efficiency rating is0.10-1.01.
 3. The turbomachine engine of claim 1, wherein the gearboxefficiency rating is 0.19-1.8.
 4. The turbomachine engine of claim 1,wherein the gear ratio is within a range of 4.5-12.0.
 5. Theturbomachine engine of claim 1, wherein the gear ratio is within a rangeof 6.0-11.0.
 6. The turbomachine engine of claim 1, wherein Q is withina range of 5-55 gallons per minute.
 7. The turbomachine engine of claim1, wherein Q is within a range of 6-36 gallons per minute.
 8. Theturbomachine engine of claim 1, wherein D is 120-216 inches.
 9. Theturbomachine engine of claim 1, wherein D is 120-192 inches.
 10. Theturbomachine engine of claim 1, wherein T is within a range of10,000-100,000 pounds force.
 11. The turbomachine engine of claim 1,wherein T is within a range of 12,000-30,000 pounds force.
 12. Theturbomachine engine of claim 1, wherein the gearbox is an epicyclicgearbox comprising a sun gear, a plurality of planet gears, and a ringgear, wherein the sun gear is the input, and wherein the ring gear isthe output.
 13. The turbomachine engine of claim 1, wherein the gearboxis an epicyclic gearbox comprising a sun gear, a plurality of planetgears, and a ring gear, wherein the sun gear is the input, wherein theplanet gears are coupled to a planet carrier, and wherein the planetcarrier is the output.
 14. The turbomachine engine of claim 1, whereinthe gearbox is a multi- stage gearbox.
 15. The turbomachine engine ofclaim 1, wherein the gearbox is a two- stage gearbox.
 16. Theturbomachine engine of claim 1, wherein the gearbox is a compoundgearbox.
 17. A turbomachine engine comprising: a fan assembly includinga single row of fan blades; a vane assembly including a plurality ofvanes disposed aft of the single row of fan blades, wherein theplurality of vanes receives bypass airflow from the single row of fanblades; a core engine including a low-pressure compressor, ahigh-pressure compressor, a combustor, a high-pressure turbine, and alow-pressure turbine; a gearbox including an input and an output,wherein the input is coupled to the low-pressure turbine and comprises afirst rotational speed, wherein the output is coupled to the fanassembly and has a second rotational speed, and wherein a gear ratio ofthe first rotational speed to the second rotational speed is within arange of 4.5-14.0; and a gearbox efficiency rating of 0.12-1.8, whereinthe gearbox efficiency rating equals Q(D{circumflex over( )}1.56/T){circumflex over ( )}1.53, wherein Q is a gearbox oil flowrate at an inlet of the gearbox measured in gallons per minute at a maxtakeoff condition, wherein D is a diameter of the fan blades measured ininches, and wherein T is a net thrust of the turbomachine enginemeasured in pounds force at the max takeoff condition.
 18. Theturbomachine engine of claim 17, wherein Q is within a range of 5-55gallons per minute, wherein D is 120-216 inches, and wherein T is withina range of 10,000-100,000 pounds force.
 19. The turbomachine engine ofclaim 17, wherein the gearbox is an epicyclic gearbox comprising a stargear configuration.
 20. The turbomachine engine of claim 17, wherein thegearbox is an epicyclic gearbox comprising a planet gear configuration.21. The turbomachine engine of claim 17, wherein the single row of fanblades comprises 8-20 fan blades, wherein the low-pressure compressorcomprises 1-8 stages, wherein the high-pressure compressor comprises8-15 stages, wherein the high-pressure turbine comprises 1-2 stages, andthe low-pressure turbine comprises 3-7 stages.
 22. A turbomachine enginecomprising: a fan assembly including a plurality of fan blades arrangedin exactly one row; a vane assembly including a plurality of vanesdisposed aft of the plurality of fan blades and radially aligned withbypass airflow directed from the plurality of fan blades to theplurality of vanes; a core engine including a low-pressure compressor, ahigh-pressure compressor, a combustor, a high-pressure turbine, and alow-pressure turbine; and a gearbox comprising a gear ratio within arange of 6.0-12.0 and a gearbox efficiency rating of 0.18-1.41.
 23. Theturbomachine engine of claim 22, wherein a gearbox oil flow rate at aninlet of the gearbox is within a range of 5-55 gallons per minute at amax takeoff condition.
 24. The turbomachine engine of claim 22, whereina diameter of the fan blades is 120-216 inches.
 25. The turbomachineengine of claim 22, wherein a net thrust of the turbomachine engine iswithin a range of 10,000-100,000 pounds force at a max takeoffcondition.
 26. A turbomachine engine comprising: a fan assemblyincluding a single row of fan blades; a vane assembly including aplurality of vanes disposed aft of and configured to receive bypassairflow from the single row of fan blades; a core engine including oneor more compressor sections and one or more turbine sections; a gearboxincluding an input and an output, wherein the input is coupled to theone or more turbine sections of the core engine and comprises a firstrotational speed, wherein the output is coupled to the fan assembly andhas a second rotational speed, and wherein a gear ratio of the firstrotational speed to the second rotational speed is within a range of4.1-13.0; and a gearbox efficiency rating of 0.19-1.8, wherein thegearbox efficiency rating equals${Q\left( \frac{D^{1.56}}{T} \right)}^{1.53},$ wherein Q is a gearboxoil flow rate at an inlet of the gearbox measured in gallons per minuteat a max takeoff condition, wherein D is a diameter of the fan bladesmeasured in inches, and wherein T is a net thrust of the turbomachineengine measured in pounds force at the max takeoff condition.
 27. Theturbomachine engine of claim 26, wherein the gearbox is an epicyclicgearbox comprising a gear ratio within a range of 4.1-9.9.
 28. Theturbomachine engine of claim 27, wherein the gearbox efficiency ratingis 0.19-1.06.
 29. The turbomachine engine of claim 26, wherein thegearbox is a non-epicyclic gearbox comprising a gear ratio within arange of 4.1-9.9.
 30. The turbomachine engine of claim 29, wherein thegearbox efficiency rating is 0.19-0.62.